CN110095990A - A kind of aircraft end direct force pulse duration modulation method - Google Patents
A kind of aircraft end direct force pulse duration modulation method Download PDFInfo
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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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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- 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
- G05B13/048—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 using a predictor
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- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D17/00—Control of torque; Control of mechanical power
- G05D17/02—Control of torque; Control of mechanical power characterised by the use of electric means
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Abstract
A kind of aircraft end direct force pulse duration modulation method of the invention, the following steps are included: analyzing existing aircraft rail control formula side-jet control mode feature, it is proposed direct force pulse width modulation control method, i.e. direct engine spout uses cross distribution, direct Thrust size is unadjustable, achievees the purpose that control the direct force duration by opening opposite valve；The side-jet control strategy of the control method, including thrust equivalence principle is proposed for the control method feature, and derives the direct engine available machine time；The control method needs to calculate residual non-uniformity, estimates miss distance according to thrust equivalence principle, devises the motor-driven estimation based on extended state observer to target；The control method is displaced correction Limits properties according to existing for thrust equivalence principle, devises the variable structure guidance law based on zero miss distance.
Description
Technical field
The present invention relates to a kind of aircraft end direct force pulse duration modulation method more particularly to a kind of direct Thrust size are permanent
Fixed, thrust time adjustable direct force momentum control technology.
Background technique
Current existing rail control formula direct force/aerodynamic force Compound Control Technique focuses primarily upon two kinds of thrust modes: a kind of
Using European Aster-15, Aster-30 as representative, direct Thrust engine is modulated by gas switch valve, and direct Thrust is big
It is small continuously adjustable；Second to be 9M96E, 9M96E2 in Russian S-400 system of defense as representative, in the mass center position of guided missile
It sets circumferentially throughout certain amount spout, by adjusting direct Thrust in control specific direction upper switch certain amount spout
There is certain lowest resolution, the company of can not achieve in direction and size, the synthesis thrust size of this regulative mode and direction
It is continuous adjustable, for convenience of describing, it is defined as the discrete adjustable control mode of direct Thrust.
The discrete adjustable control mode of direct Thrust must be circumferentially arranged to improve the resolution ratio of thrust adjusting along body
A certain number of spouts, and the same time can only open specified quantity spout and provide to use and overload.Since multiple spouts are synthesizing
There are power counteractings on perpendicular to the direction for needing thrust when thrust, therefore can be carried fuel by body and be limited, and can provide most
High thrust is not high.The continuously adjustable control method of direct Thrust can provide in real time need to be with overload, ultimate challenge fast-response,
The high-temperature fuel gas regulating valve of high reliability.Although valve mechanism can be simplified by reducing fuel gas temperature, reliability and craftsmanship are improved,
The energy quality characteristic of engine is also reduced simultaneously.Therefore both main traditional rail control formula direct Thrust sides
All there are respective advantage and disadvantage in formula, and these advantage and disadvantage can make up mutually.
Summary of the invention
The present invention provides a kind of aircraft end direct force on the basis of comprehensive analysis existing two kinds of rails control thrust mode
Pulse duration modulation method.
In aircraft end direct force pulse duration modulation method provided by the invention, the spout of direct force engine is located at guided missile
Plane where mass center, and use cross distribution, the plane perpendicular to guided missile central axis, what direct force engine provided
Direct force size is fixed, and controls direct force end momentum by control direct force engine operating duration, step includes:
Step 1: the correction displacement that design momentum control generates is equal to the equivalence principle for estimating miss distance, determines and directly send out
Motivation opens strategy, calculates the direct force duration；
Step 2:, using mesh Equation of Relative Motion with Small is played, being estimated based on extended state observer according to the equivalence principle
Off-target value is estimated in target motion information, calculating；
Step 3: designing the sliding-mode guidance based on zero miss distance according to the equivalence principle.
Further, the step 1 includes:
Step 1.1 when meeting compressor start up condition, in fore-and-aft plane, estimates miss distance are as follows:
Wherein, ZEM indicates that target does not do motor-driven Zero effort miss distance, TgoIndicate residual non-uniformity, RtIndicate that t moment is opened
Bullet mesh relative distance when machine, VRtIndicate bullet mesh relative velocity when t moment booting, aRyIndicate that mesh is played in fore-and-aft plane to be added relatively
Component of the speed perpendicular to relative velocity；TgoFor be switched on moment residual non-uniformity, by formulaIt determines, wherein VRIt is guided missile mesh for missile target relative velocity, μ
Mark the angle of relative velocity and direction of visual lines；
Step 1.2, direct force displacement correction amount Δ S, which are equal to, estimates miss distance,nt
For the overload size that direct force generates, the direct force engine continuous time, Δ t was obtained by the following method:
(1) whenWhen, Δ t=Tgo；
(2) whenWhen, Δ t is determined by following formula:
After the direct force available machine time reaches Δ t, reach closing direct force effect by opening opposite spout.
Further, the step 2 includes:
Step 2.1 establishes original equation, to simplify the process, only considers to play mesh Equation of Relative Motion with Small in fore-and-aft plane
In formula,For line of sight rate,Be the angle of sight for LOS angle acceleration, q, R be play mesh relative distance,For bullet
Mesh size is to relative velocity, VtFor target velocity,For target velocity inclination angle speed,For missile velocity inclination angle speed, θtFor mesh
Mark speed inclination angle, θmFor missile velocity inclination angle；
Step 2.2 enablesAnd assumeIt builds
Vertical state equation:
Step 2.3, according to ESO method, establish extended state observer:Pass through observation
Value z2Obtain target tangential acceleration
Step 2.4 establishes bullet mesh Equation of Relative Motion with Small
Extended state observer is designed, is obtained
To target normal acceleration
Step 2.5, according to target normal acceleration and target tangential acceleration, obtain the aimed acceleration in fore-and-aft plane
In component estimated value on sight normal
Step 2.6, the similarly aimed acceleration in available transverse plane are in component estimated value on sight normal
Further, the step 3 includes:
Step 3.1 designs the overload instruction containing structure changes item in fore-and-aft plane
Wherein, NvFor navigation ratio, nx1For axial acceleration, the ε measured on bulletbmFor Seeker's light axis corner, aymFor guided missile plus
Speed is in component, ε on sight normal are proportionality coefficient in structure changes item, the diverter surface of s variable structure guidance law, g are that gravity accelerates
Degree；
Step 3.2 designs the overload instruction containing structure changes item in transverse plane
Wherein, NhFor navigation ratio, ψvmTrajectory deflection angle, qβFor the angle of sight in transverse plane.
The invention has the advantages that
The direct force engine of guided missile rail control presses cross distribution, and direct Thrust size is unadjustable, and switch valve controls direct force
Opening time, the control mode adjusted using momentum.This thrust mode, the simple engineering of engine physical structure are easily realized, are produced
Raw thrust is stablized, and since there is no thrust synthesis process, therefore the maximum thrust that can be generated is big.
Above-mentioned aircraft end direct force pulse duration modulation method, wherein precise tracking is directly sprayed using cross arrangement
Pipe can provide direct force in 8 directions, and after guidance end direct force is opened, thrust constant magnitude is unadjustable, passes through opening
Opposite valve reaches closing effect, controls thrust duration, and a direction is only switchable primary, using momentum regulative mode,
Achieve the purpose that endgame correction miss distance.With equivalent principle is displaced, by estimating residual non-uniformity, target maneuver size,
It estimates miss distance, then is equal to by corrected range and estimates miss distance, calculate the direct force engine continuous time；The control method
Since there are critical distances for displacement correction, therefore the variable structure guidance law based on zero miss distance is designed, guarantees that guided missile is sent out in direct force
Motivation is switched on the moment, estimates miss distance in displacement correction limit range.
Detailed description of the invention
Aircraft end direct force pulse duration modulation method of the invention is provided by examples and drawings below.
Fig. 1 is that rail control is started and illustraton of model in the embodiment of the present invention.
Fig. 2 is that the embodiment of the present invention is hit by a bullet mesh relative motion relation figure, in fore-and-aft plane.
Specific embodiment
Guided missile direct force impulse control method of the invention is described in further detail below with reference to FIG. 1 to FIG. 2.
For convenience of description, below in fore-and-aft plane.
The aircraft end direct force pulse duration modulation method of the embodiment of the present invention the following steps are included:
Step 1, the bullet mesh relative information that missile-borne computer is provided according to radar seeker, including line of sight rateBullet
Mesh relative distance R plays mesh size to relative velocityAnd the Missile Motion status information resolved according to strapdown, utilize bullet
Mesh Equation of Relative Motion with Small estimating target motion information:
V in formulatFor target velocity,For target velocity inclination angle speed,For missile velocity inclination angle speed, θtFor target speed
Spend inclination angle, θmFor missile velocity inclination angle；Target tangential acceleration information is unknown, and needs are estimated.
Therefore it can enableAnd assumeThen may be used
Establish state equation:
According to ESO method, extended state observer is established:
Pass through observation z2Then available target tangential acceleration information.
Similarly by bullet mesh Equation of Relative Motion with Small:
Extended state observer, available target normal direction acceleration information are designed, synthesis can obtain target in fore-and-aft plane
Acceleration information.It can get aimed acceleration estimated value in laterally and longitudinally plane using the method for ESO.
Step 2, by taking fore-and-aft plane as an example, the overload instruction of design item containing structure changes
It enablesTake state variableInterference volume isIndicate state observer estimate error, control amountTherefore play mesh relative motion side
Journey can turn to:
It needs to take and estimates miss distance as sliding formwork diverter surface:
Design Reaching Law are as follows:
It is required according to variable structure guidance law reaching conditionWherein δ is arbitrarily small positive number, can be obtained
NvFor navigation ratio, nx1For axial acceleration, the ε measured on bulletbmFor Seeker's light axis corner, aymFor guided missile acceleration
In component, ε on sight normal are proportionality coefficient in structure changes item, the diverter surface of s variable structure guidance law, g are acceleration of gravity.
Overload instruction in transverse plane, which can similarly be obtained, is
NhFor navigation ratio, ψvmTrajectory deflection angle, qβFor the angle of sight in transverse plane.
It step 3, is the variable structure guidance law based on zero miss distance according to the Terminal Guidance Laws of the overload instruction design of step 2,
Purpose is to ensure that direct force start-up time estimates miss distance and is less than the displacement correction limit.Judge at the terminal guidance section momentDiverging
Situation, when meeting compressor start up condition, miss distance is estimated in calculating at this time:
Wherein, ZEM indicates that target does not do motor-driven Zero effort miss distance, TgoIndicate residual non-uniformity, RtIndicate that t moment is opened
Bullet mesh relative distance when machine, VRtIndicate bullet mesh relative velocity when t moment booting；aRyIt indicates to play mesh relative acceleration, utilize
IMU resolving value is obtained with state observation estimated value；TgoFor be switched on moment residual non-uniformity, by formula:
Step 4, it under the conditions of obtaining estimating miss distance ZEM, is displaced and is rectified a deviation by direct force:
ntFor the overload size that direct force generates, then direct force engine continuous time Δ t is obtained by the following method:
(1) whenWhen, Δ t=Tgo；
(2) whenWhen, Δ t is determined by following formula:
After the direct force available machine time reaches Δ t, reach closing direct force effect by opening opposite spout.It can from above formula
It to obtain estimating miss distance out and need to estimate target maneuver information；Direct force displacement correction amount has maximum distance simultaneously, therefore needs
Design the guidance law met.This aircraft end direct force pulse duration modulation method, the physics knot of engine and switch valve
Structure is simply easily achieved, and engine operation and the thrust of offer are stablized, and maximum thrust is big, great reality and theoretical research
Meaning.Thrust equivalence principle that the present invention designs, the variable structure guidance law based on zero miss distance all have other types guided missile
Reference.
Claims (4)
1. a kind of aircraft end direct force pulse duration modulation method, which is characterized in that the spout of direct force engine is located at guided missile
Plane where mass center, and use cross distribution, the plane perpendicular to guided missile central axis, what direct force engine provided
Direct force size is fixed, and controls direct force end momentum by control direct force engine operating duration, step includes:
Step 1: the correction displacement that design momentum control generates is equal to the equivalence principle for estimating miss distance, direct engine is determined
Strategy is opened, the direct force duration is calculated；
Step 2:, using mesh Equation of Relative Motion with Small is played, estimating target based on extended state observer according to the equivalence principle
Off-target value is estimated in motion information, calculating；
Step 3: designing the sliding-mode guidance based on zero miss distance according to the equivalence principle.
2. a kind of aircraft end direct force pulse duration modulation method according to claim 1, which is characterized in that the step
One includes:
Step 1.1 when meeting compressor start up condition, in fore-and-aft plane, estimates miss distance are as follows:
Wherein, ZEM indicates that target does not do motor-driven Zero effort miss distance, TgoIndicate residual non-uniformity, RtWhen indicating t moment booting
Bullet mesh relative distance, VRtIndicate bullet mesh relative velocity when t moment booting, aRyIt indicates to play mesh relative acceleration in fore-and-aft plane
Perpendicular to the component of relative velocity；TgoFor be switched on moment residual non-uniformity, by formulaIt determines, wherein VRIt is guided missile mesh for missile target relative velocity, μ
Mark the angle of relative velocity and direction of visual lines；
Step 1.2, direct force displacement correction amount Δ S, which are equal to, estimates miss distance,ntIt is direct
The overload size that power generates, the direct force engine continuous time, Δ t was obtained by the following method:
(1) whenWhen, Δ t=Tgo；
(2) whenWhen, Δ t is determined by following formula:Directly
After the relay available machine time reaches Δ t, reach closing direct force effect by opening opposite spout.
3. a kind of aircraft end direct force pulse duration modulation method according to claim 2, which is characterized in that the step
Two include:
Step 2.1 establishes original equation, to simplify the process, only considers to play mesh Equation of Relative Motion with Small in fore-and-aft plane:
In formula,For line of sight rate,Be the angle of sight for LOS angle acceleration, q, R be play mesh relative distance,To play mesh size
To relative velocity, VtFor target velocity,For target velocity inclination angle speed,For missile velocity inclination angle speed, θtFor target speed
Spend inclination angle, θmFor missile velocity inclination angle；
Step 2.2 enablesAnd assumeEstablish shape
State equation:
Step 2.3, according to ESO method, establish extended state observer:Pass through observation z2
Obtain target tangential acceleration
Step 2.4 establishes bullet mesh Equation of Relative Motion with Small
Extended state observer is designed, mesh is obtained
Mark normal acceleration
Step 2.5, according to target normal acceleration and target tangential acceleration, obtain the aimed acceleration in fore-and-aft plane in bullet
Mesh line normal component estimated value
Step 2.6, the similarly aimed acceleration in available transverse plane are in component estimated value on bullet mesh line normal
4. a kind of aircraft end direct force pulse duration modulation method according to claim 3, which is characterized in that the step
Three include:
Step 3.1 designs the overload instruction containing structure changes item in fore-and-aft plane
Wherein, NvFor navigation ratio, nx1For axial acceleration, the ε measured on bulletbmFor Seeker's light axis corner, aymFor guided missile acceleration
In play the component on mesh line normal, ε is proportionality coefficient in structure changes item, the diverter surface of s variable structure guidance law, g are that gravity adds
Speed；
Step 3.2 designs the overload instruction containing structure changes item in transverse plane
Wherein, NhFor navigation ratio, ψvmTrajectory deflection angle, qβFor the angle of sight, a in transverse planezmIndicate that guided missile accelerates in transverse plane
It spends in the component on bullet mesh line normal.
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110597056A (en) * | 2019-08-16 | 2019-12-20 | 南京理工大学 | Large closed-loop calibration control method for antiaircraft gun fire control system |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN110597056A (en) * | 2019-08-16 | 2019-12-20 | 南京理工大学 | Large closed-loop calibration control method for antiaircraft gun fire control system |
| CN110597056B (en) * | 2019-08-16 | 2022-06-28 | 南京理工大学 | Large closed-loop calibration control method for antiaircraft gun fire control system |
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