CN116700306B - Integrated guidance control method for strapdown guided aircraft - Google Patents

Integrated guidance control method for strapdown guided aircraft Download PDF

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CN116700306B
CN116700306B CN202310646328.XA CN202310646328A CN116700306B CN 116700306 B CN116700306 B CN 116700306B CN 202310646328 A CN202310646328 A CN 202310646328A CN 116700306 B CN116700306 B CN 116700306B
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aircraft
angle
sight
state quantity
angular velocity
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CN116700306A (en
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王伟
刘佳琪
林时尧
王少龙
耿宝魁
张宏岩
于之晨
李俊辉
李成洋
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Northwest Industrial Group Co ltd
Beijing Institute of Technology BIT
Ordnance Science and Research Academy of China
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Northwest Industrial Group Co ltd
Beijing Institute of Technology BIT
Ordnance Science and Research Academy of China
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Abstract

S1, setting a control system model, and describing the relation among the sight angle, the sight angular speed, the aircraft posture, the speed and the aircraft control quantity of the aircraft through the control system model; s2, according to a control system model, acquiring an aircraft sight angle, a sight angle speed, an aircraft gesture and a speed of the aircraft at the current moment and an aircraft control quantity at the last moment, acquiring a current control signal, and controlling the aircraft to deflect according to the control signal by the aircraft. The integrated guidance control method for the strapdown guided aircraft disclosed by the invention can ensure that the target is always in the capture range of the guide head, and has higher control precision.

Description

Integrated guidance control method for strapdown guided aircraft
Technical Field
The invention relates to an integrated guidance control method for a strapdown guided aircraft, and belongs to the field of aircraft guidance.
Background
In the process of flying to a target, the conventional strapdown guidance aircraft generally leads the guidance head to point to the target in a passive adjustment mode, and when the target is maneuvering greatly or the aircraft shakes in the flight process, after the strapdown guidance head with a narrow sight is in the target escape visual field, the guidance head is difficult to quickly align by directly controlling the aircraft.
The integrated guidance control system design facilitates efficient control of the aircraft. For an integrated guidance control system of an aircraft designed by a traditional back-stepping method, the next derivative is obtained by directly solving a virtual control quantity in the design process, but when the virtual control quantity is rapidly changed, the solved derivative variable rapidly expands and changes due to direct differentiation, so that the control process is influenced.
Therefore, there is a need to propose an integrated guidance control method for strapdown guided aircraft to solve the above problems.
Disclosure of Invention
In order to overcome the above problems, the present inventors have conducted intensive studies and devised an integrated guidance control method for a strapdown guided vehicle, comprising the steps of:
s1, setting a control system model, and describing the relation among the sight angle, the sight angular speed, the attitude and the speed of the aircraft and the control quantity of the aircraft through the control system model;
s2, according to a control system model, acquiring an aircraft sight angle, a sight angle speed, an aircraft gesture and a speed of the aircraft at the current moment and an aircraft control quantity at the last moment, acquiring a current control signal, and controlling the aircraft to deflect according to the control signal by the aircraft.
In a preferred embodiment, in S1, a control system model is constructed using the aircraft viewing angle, viewing angular velocity, aircraft attitude, and angular velocity as model state amounts and the control amount of the aircraft as an input amount;
wherein the line of sight angle state quantity is expressed as，/>，/>For the angle between the axis direction of the aircraft and the target in the direction of the inclination of the line of sight +.>The included angle between the axis direction of the aircraft and the target in the direction of the deflection angle of the sight;
the state quantity of the angular velocity of the line of sight is expressed as，/>，/>Represents the inclination of the line of sight with respect to the target with the origin of the aircraft, < >>Representing a line of sight offset angle with respect to the target with the aircraft as an origin;
the attitude state quantity is expressed as，/>，/>For the angle of attack of the aircraft, +.>Sideslip angle for aircraft, +.>Is the roll angle of the aircraft;
the angular velocity state quantity is expressed as，/>，/>For pitch rate of aircraft,/->For yaw rate of the aircraft, +.>Is the roll angle speed of the aircraft;
the aircraft control quantity is expressed as，/>，/>For controlling the equivalent rudder deflection angle of the pitch of the aircraft, < >>Equivalent rudder deflection angle for controlling the yaw of an aircraft, < >>For controlling the equivalent rudder deflection angle of the aircraft roll.
In a preferred embodiment, in S1, the control system model is expressed as:
，
wherein,
，
，
，
，
，
，
，
wherein,representing pitch angle, < >>Represents the yaw angle of the aircraft, and has:
，
，
representing the relative distance between the aircraft and the target, < +.>Representing dynamic pressure of aircraft->，/>For air density->Is the speed of the aircraft, +.>Representing thrust of aircraft, ++>Representing aircraft mass, +.>Representing aircraft reference area, < >>Representing the characteristic length of an aircraft, < >>Representing the lift coefficient; />Representing the lateral force coefficient; />Representing the lateral static derivative induced by the angle of attack;representing the lateral static derivative induced by the sideslip angle; />Representing yaw stationarity derivative,/->Representing the derivative of the static pitch stability,indicating aileron steering efficiency, +.>Indicating rudder steering efficiency,/->Indicating elevator operating efficiency,/->Representing the moment of inertia of the aircraft about the longitudinal axis, < >>Representing the moment of inertia of the aircraft about the vertical axis, +.>Representing the moment of inertia of the aircraft about the transverse axis.
In a preferred embodiment, S2 comprises the sub-steps of:
s21, setting a first error face, taking a desired value of the state quantity of the angular velocity of the sight line as virtual input, enabling the angle of the sight line to be 0, and obtaining the change rate of the desired value of the state quantity of the angular velocity of the sight line through a first improved sliding mode differentiator;
s22, setting a second error surface to enable the expected value of the state quantity of the line of sight angular velocity and the state quantity difference value of the line of sight angular velocity to be 0, and obtaining the change rate of the expected value of the state quantity of the attitude through a second improved sliding mode differentiator;
s23, setting a third error face to enable the expected value of the attitude state quantity and the difference value of the attitude state quantity to be 0, and obtaining the change rate of the expected value of the angular velocity state quantity through a third improved sliding mode differentiator;
s24, setting a fourth error face to enable the expected value of the angular velocity state quantity and the difference value of the angular velocity state quantity to be 0, and obtaining the control quantity of the aircraft.
In a preferred embodiment, in S21, the first error plane is，
Setting the expected value of the state quantity of the line of sight angular velocity to be solved as follows:thereby obtaining a rate of change of the desired value of the state quantity of the angular velocity of view through said first modified sliding mode differentiator, wherein +.>Is constant.
In a preferred embodiment, the first modified sliding mode differentiator is represented as:
，
wherein,representation->Estimate of +.>To estimate->A set higher-order variable, and has +.>。
In a preferred embodiment, in S22, the second error plane is:，
setting the expected value of the attitude state quantity to be solved as follows:
，
，
and further obtaining a rate of change of the expected value of the attitude state quantity by a second modified sliding mode differentiator, wherein,is constant.
In a preferred embodiment, the second modified sliding mode differentiator is represented as:
，
wherein,representation->Estimate of +.>To estimate->Designed higher-order variables and have +.>。
In a preferred embodiment, in S23, the third error plane is:，
setting the expected value of the state quantity of the angular velocity to be solved as follows:
，
wherein,is constant.
In a preferred embodiment, in S24, the fourth error plane is:，
the fourth error face is derived to obtain:
，
the aircraft control amounts obtained were:
，
wherein,is constant.
The invention has the beneficial effects that:
(1) According to the integrated guidance control method for the strapdown guided aircraft, provided by the invention, the target can be always in the capture range of the guide head;
(2) According to the integrated guidance control method for the strapdown guided aircraft, provided by the invention, guidance and control effects can be realized simultaneously by designing the input quantity;
(3) According to the integrated guidance control method for the strapdown guided aircraft, provided by the invention, the available precision can be ensured in the process of obtaining the approximate derivative of the design quantity.
Drawings
FIG. 1 illustrates a flow diagram of an integrated guidance control method for a strapdown guided vehicle in accordance with a preferred embodiment of the present invention;
FIG. 2 shows simulation results of the change in position of a target in the strapdown field of view in example 1;
fig. 3 shows simulation results of example 1 and comparative example 1 for target tracking.
Detailed Description
The invention is further described in detail below by means of the figures and examples. The features and advantages of the present invention will become more apparent from the description.
The word "exemplary" is used herein to mean "serving as an example, embodiment, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Although various aspects of the embodiments are illustrated in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated.
The integrated guidance control system design facilitates efficient control of the aircraft. The integrated guidance control refers to that after control feedback information of the speed, the position, the attitude angle, the strapdown seeker sight angle and the deflection mechanism of the aircraft is obtained, integrated processing is carried out on the information through an onboard processor, and guidance control instructions are directly output.
The integrated guidance control method for the strapdown guided aircraft provided by the invention comprises the following steps of:
s1, setting a control system model, and describing the relation among the sight angle, the sight angular speed, the attitude, the angular speed and the control quantity of the aircraft through the control system model;
s2, acquiring the sight angle, the sight angular speed, the attitude and the angular speed of the aircraft at the current moment and the control quantity of the aircraft at the previous moment according to a control system model, obtaining a current control signal, and controlling the aircraft to deflect according to the control signal by the aircraft.
According to the invention, in S1, the sight angle, the sight angular velocity, the attitude and the angular velocity of the aircraft are respectively taken as model state quantities, and the control quantity of the aircraft is taken as an input quantity to construct a control system model;
wherein the line of sight angle state quantity is expressed asFor describing the angle of view of the aircraft with respect to the target, < >>，/>For the angle between the axis direction of the aircraft and the target in the direction of the inclination of the line of sight +.>The included angle between the axis direction of the aircraft and the target in the direction of the deflection angle of the sight;
the state quantity of the angular velocity of the line of sight is expressed asFor describing the angular velocity of view of an aircraft relative to a target,/->，/>Represents the inclination of the line of sight with respect to the target with the origin of the aircraft, < >>Representing a line of sight offset angle with respect to the target with the aircraft as an origin;
the attitude state quantity is expressed asFor describing the attitude of an aircraft, +.>，/>For the angle of attack of the aircraft, +.>Sideslip angle for aircraft, +.>Is the roll angle of the aircraft;
the angular velocity state quantity is expressed asFor describing the flight rotational speed of an aircraft, < +.>，/>For pitch rate of aircraft,/->For yaw rate of the aircraft, +.>For flyingThe roll angle speed of the device;
the aircraft control quantity is expressed as，/>，/>For controlling the equivalent rudder deflection angle of the pitch of the aircraft, < >>Equivalent rudder deflection angle for controlling the yaw of an aircraft, < >>For controlling the equivalent rudder deflection angle of the aircraft roll.
Further, in S1, the control system model is expressed as:
，
，
，
，
，
，
，
，
wherein,representing pitch angle, < >>Represents the yaw angle of the aircraft, and has:
，
，
representing the relative distance between the aircraft and the target, < +.>Representing dynamic pressure of aircraft->，/>In order to achieve an air density of the air,is the speed of the aircraft, +.>Representing thrust of aircraft, ++>Representing aircraft mass, +.>Representation ofAircraft reference area->Representing the characteristic length of an aircraft, < >>Representing the lift coefficient; />Representing the lateral force coefficient; />Representing the lateral static derivative induced by the angle of attack; />Representing the lateral static derivative induced by the sideslip angle; />Representing yaw stationarity derivative,/->Representing the static derivative of pitch>Indicating aileron steering efficiency, +.>Indicating rudder steering efficiency,/->Indicating elevator operating efficiency,/->Representing the moment of inertia of the aircraft about the longitudinal axis, < >>Representing the moment of inertia of the aircraft about the vertical axis, +.>Representing the moment of inertia of the aircraft about the transverse axis.
The model is simplified while the main dynamics and kinematic processes of the aircraft can be reflected, the control efficiency is improved, and meanwhile, the view angle constraint item of the aircraft is introduced, so that the view angle range of the aircraft can be constrained.
In S2, the current control signal is obtained by a back-step control method.
In the invention, the virtual control quantity is improved by the sliding mode differentiator, so that the extremely-variable situation of the required conductivity can be avoided, and the improved sliding mode differentiator can obtain the virtual control quantity derivative with higher precision compared with the traditional sliding mode differentiator.
S2 comprises the following substeps:
s21, setting a first error face, taking a desired value of the state quantity of the line of sight angular velocity as a virtual input, enabling the line of sight angle to be 0, namely controlling the included angle between the axis of the aircraft and the connecting line of the mass center of the aircraft and the target to be converged to 0, enabling the aircraft to be aligned to the target, and obtaining the change rate of the desired value of the state quantity of the line of sight angular velocity through a first improved sliding mode differentiator;
s22, setting a second error surface to enable the expected value of the state quantity of the line of sight angular velocity and the state quantity difference value of the line of sight angular velocity to be 0, and obtaining the change rate of the expected value of the state quantity of the attitude through a second improved sliding mode differentiator;
s23, setting a third error face to enable the expected value of the attitude state quantity and the difference value of the attitude state quantity to be 0, and obtaining the change rate of the expected value of the angular velocity state quantity through a third improved sliding mode differentiator;
s24, setting a fourth error face to enable the expected value of the angular velocity state quantity and the difference value of the angular velocity state quantity to be 0, and obtaining the control quantity of the aircraft.
Specifically, in S21, the first error plane is，
Setting the expected value of the state quantity of the line of sight angular velocity to be solved as follows:thereby obtaining a rate of change of the desired value of the state quantity of the angular velocity of view through said first improved sliding mode differentiator, wherein +.>Is constant.
The expected value setting mode can realize the control quantity required by the step through only one parameter, and has simple structure and easy realization.
Further preferably, the first modified sliding mode differentiator is represented as:
，
wherein,representation->Estimate of +.>To estimate->A set higher-order variable, and has +.>。
An improvement in the accuracy of the estimation of the desired signal is achieved by adding a non-linear term in the first modified sliding mode differentiator.
In a preferred embodiment, in S22, the second error plane is:，
setting the expected value of the attitude state quantity to be solved as follows:
，
，
and further obtaining a rate of change of the expected value of the attitude state quantity by a second modified sliding mode differentiator, wherein,is constant.
Likewise, the above-described desired value setting method realizes a desired control amount by introducing only one parameter, and facilitates the realization of control by simplifying the structure.
Further preferably, the second modified sliding mode differentiator is represented as:
，
wherein,representation->Estimate of +.>To estimate->Designed higher-order variables and have +.>。
Likewise, by adding a nonlinear term in the second modified sliding mode differentiator, an improvement in the accuracy of the estimation of the desired signal is achieved.
In a preferred embodiment, in S23, the third error plane is:，
the method can be obtained by deriving the following steps:
。
in a preferred embodiment, the desired value of the state quantity of angular velocity to be solved is set as:
，
wherein,is constant.
Likewise, the above-described desired value setting method realizes a desired control amount by introducing only one parameter, and facilitates the realization of control by simplifying the structure.
Further preferably, the third modified sliding mode differentiator is represented as:
，
wherein,representation->Estimate of +.>To estimate->Designed higher-order variables and have +.>。
Likewise, by adding a nonlinear term in the second improved sliding mode differentiator, the desired signal is achievedThe estimation accuracy of the number is improved. In a preferred embodiment, in S24, the fourth error plane is:。
further, deriving the fourth error plane may result in:
，
the aircraft control values thus obtained were:
，
wherein,is constant.
According to a preferred embodiment of the inventionThe value range of (2) is 1-10, (-)>The range of the value of (2) is 5-15, < + >>The range of the value of (2) is 15-25 #>The value of (2) is 25-35, more preferably, ">。
Examples
Example 1
Performing simulation experiments, wherein the aircraft tracks a target, and the target is accelerated in a serpentine shape as followsMovement:
，
the aircraft tracks the target using the following steps:
s1, setting a control system model, and describing the relation among the sight angle, the sight angular speed, the attitude, the angular speed and the control quantity of the aircraft through the control system model;
s2, acquiring the sight angle, the sight angular speed, the attitude and the angular speed of the aircraft at the current moment and the control quantity of the aircraft at the previous moment according to a control system model, obtaining a current control signal, and controlling the aircraft to deflect according to the control signal by the aircraft.
In S1, the control system model is expressed as:
，
，
，
，
，
，
，
，
wherein:
，
。
in S2, the current control signal is obtained by a back-step control method, comprising the sub-steps of:
s21, setting a first error face, taking a desired value of the state quantity of the angular velocity of the sight line as virtual input, enabling the angle of the sight line to be 0, and obtaining the change rate of the desired value of the state quantity of the angular velocity of the sight line through a first improved sliding mode differentiator;
s22, setting a second error surface to enable the expected value of the state quantity of the line of sight angular velocity and the state quantity difference value of the line of sight angular velocity to be 0, and obtaining the change rate of the expected value of the state quantity of the attitude through a second improved sliding mode differentiator;
s23, setting a third error face to enable the expected value of the attitude state quantity and the difference value of the attitude state quantity to be 0, and obtaining the change rate of the expected value of the speed state quantity through a third improved sliding mode differentiator;
s24, setting a fourth error face to enable the expected value of the speed state quantity and the speed state quantity difference value to be 0, and obtaining the control quantity of the aircraft.
Specifically, in S21, the first error plane is，
Setting the expected value of the state quantity of the line of sight angular velocity to be solved as follows:wherein (1)>5.
The first modified sliding mode differentiator is denoted as:
，
wherein,。
in S22, the second error plane is:，
setting the expected value of the attitude state quantity to be solved as follows:
，
，
wherein,10.
The second modified sliding mode differentiator is denoted as:
，
wherein,。
in S23, the third error plane is:，
the method can be obtained by deriving the following steps:
。
setting the expected value of the state quantity of the angular velocity to be solved as follows:
，
wherein,20.
The third modified sliding mode differentiator is denoted as:
，
wherein,。
in S24, the fourth error plane is:，
the fourth error face is derived to obtain:
。
the aircraft control values thus obtained were:
，
wherein,30.
In tracking the target, the position of the target in the strapdown field of view changes as shown in fig. 2, it can be seen that the aircraft can always lock the target within the effective field of view of the seeker, thereby achieving a final hit.
Comparative example 1
The same simulation experiment as in example 1 was performed, except that a first-order low-pass filter was used instead of the modified sliding-mode differentiator in step S2, to obtain a simulation result.
The signal tracking results obtained in the embodiment 1 and the comparative example 1 are shown in fig. 3, and it can be seen from the graph that the embodiment 1 has a faster response speed and tracking accuracy for tracking the target compared with the comparative example 1, and especially, at the initial time of finding the target, the tracking of the target can be rapidly realized, and the target loss is avoided.
In the description of the present invention, it should be noted that the positional or positional relationship indicated by the terms such as "upper", "lower", "inner", "outer", "front", "rear", etc. are based on the positional or positional relationship in the operation state of the present invention, are merely for convenience of describing the present invention and simplifying the description, and do not indicate or imply that the apparatus or elements referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus should not be construed as limiting the present invention. Furthermore, the terms "first," "second," "third," "fourth," and the like are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
The invention has been described above in connection with preferred embodiments, which are, however, exemplary only and for illustrative purposes. On this basis, the invention can be subjected to various substitutions and improvements, and all fall within the protection scope of the invention.

Claims (3)

1. An integrated guidance control method for a strapdown guided vehicle, comprising the steps of:
s1, setting a control system model, and describing the relation among the sight angle, the sight angular speed, the attitude and the speed of the aircraft and the control quantity of the aircraft through the control system model;
s2, acquiring an aircraft sight angle, a sight angle speed, an aircraft gesture and a speed of the aircraft at the current moment and an aircraft control quantity at the previous moment according to a control system model, acquiring a current control signal, and controlling the aircraft to deflect according to the control signal by the aircraft;
in S1, taking an aircraft sight angle, a sight angular velocity, an aircraft gesture and an angular velocity as model state quantities respectively, and taking a control quantity of an aircraft as an input quantity to construct a control system model;
wherein the state quantity of the sight angle is expressed as x 0 ，x 0 ＝[ε B η B ] T ，ε B Is the included angle eta between the axis direction of the aircraft and the target in the direction of the inclination angle of the sight line B The included angle between the axis direction of the aircraft and the target in the direction of the deflection angle of the sight;
the line of sight angular velocity state quantity is expressed as x 1 ，Epsilon represents the inclination angle of the sight line with respect to the target by taking the aircraft as the origin, and eta represents the deflection angle of the sight line with respect to the target by taking the aircraft as the origin;
the attitude state quantity is expressed as x 2 ，x 2 ＝[α β γ] T Alpha is the attack angle of the aircraft, beta is the sideslip angle of the aircraft, and gamma is the roll angle of the aircraft;
the angular velocity state quantity is expressed as x 3 ，x 3 ＝[ω z ω y ω x ] T ，ω z For pitch rate, omega of aircraft y For yaw rate, ω of an aircraft x Is the roll angle speed of the aircraft;
the aircraft control quantity is expressed as u, u= [ delta ] z δ y δ x ] T ，δ z For controlling the equivalent rudder deflection angle, delta of the pitch of an aircraft y Equivalent rudder deflection angle delta for controlling yaw of aircraft x An equivalent rudder deflection angle for controlling the roll of the aircraft;
in S1, the control system model is expressed as:
wherein,
where θ represents the aircraft pitch angle, ψ represents the aircraft yaw angle, and there are:
r represents the relative distance between the aircraft and the target, q represents the dynamic pressure of the aircraft, q=0.5 ρv 2 ρ is the air density, V is the aircraft speed, P is the aircraft thrust, m is the aircraft mass, S is the aircraft reference area, L is the aircraft characteristic length,representing the lift coefficient; />Representing the lateral force coefficient; />Representing the lateral static derivative induced by the angle of attack; />Representing the lateral static derivative induced by the sideslip angle; />Representing yaw stationarity derivative,/->Representing the static derivative of pitch>Indicating aileron steering efficiency, +.>Indicating rudder steering efficiency,/->Indicating elevator operating efficiency, J x Representing moment of inertia of the aircraft about a longitudinal axis, J y Representing moment of inertia of the aircraft about a vertical axis, J z Representing the moment of inertia of the aircraft about the transverse axis;
s2 comprises the following substeps:
s21, setting a first error face, taking a desired value of the state quantity of the angular velocity of the sight line as virtual input, enabling the angle of the sight line to be 0, and obtaining the change rate of the desired value of the state quantity of the angular velocity of the sight line through a first improved sliding mode differentiator;
s22, setting a second error surface to enable the expected value of the state quantity of the line of sight angular velocity and the state quantity difference value of the line of sight angular velocity to be 0, and obtaining the change rate of the expected value of the state quantity of the attitude through a second improved sliding mode differentiator;
s23, setting a third error face to enable the expected value of the attitude state quantity and the difference value of the attitude state quantity to be 0, and obtaining the change rate of the expected value of the angular velocity state quantity through a third improved sliding mode differentiator;
s24, setting a fourth error surface to enable the expected value of the angular velocity state quantity and the difference value of the angular velocity state quantity to be 0, and obtaining the control quantity of the aircraft;
in S21, the first error face is S 0 ＝x 0 -0，
Setting the expected value of the state quantity of the line of sight angular velocity to be solved as follows: x is x 1d ＝-K 0 s 0 -f 0 Thereby obtaining a rate of change of the desired value of the line of sight angular velocity state quantity by the first modified sliding mode differentiator, wherein K 0 Is a constant;
the first modified sliding mode differentiator is denoted as:
wherein z is 11 Represents x 1d Estimated value of z of (2) 12 To estimateA set higher-order variable, and has +.>
In S22, the second error plane is: s is(s) 1 ＝x 1 -x 1d
Setting the expected value of the attitude state quantity to be solved as follows:
and then the attitude state quantity is obtained through a second improved sliding mode differentiatorRate of change of desired value, where K 1 Is a constant;
in S23, the third error plane is: s is(s) 2 ＝x 2 -x 2d
Setting the expected value of the state quantity of the angular velocity to be solved as follows:
wherein K is 2 Is a constant;
in S24, the fourth error plane is: s is(s) 3 ＝x 3 -x 3d 。
2. The integrated guidance control method for a strapdown guided vehicle of claim 1, wherein,
the second modified sliding mode differentiator is denoted as:
wherein z is 21 Represents x 2d Estimated value of z of (2) 22 To estimateDesigned higher-order variables and have +.>
3. The integrated guidance control method for a strapdown guided vehicle of claim 2, wherein,
in S24 the process proceeds to the step of,
the fourth error face is derived to obtain:
the aircraft control amounts obtained were:
wherein K is 3 Is constant.
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