CN116483109B - Sliding mode control-based integrated aircraft guidance control method and system - Google Patents
Sliding mode control-based integrated aircraft guidance control method and system Download PDFInfo
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Abstract
The application relates to an integrated method and system for aircraft guidance control based on sliding mode control, belongs to the technical field of guidance control, and solves the problem that the existing sliding mode control approaches slowly. Comprising the following steps: converting the established guidance control integrated model into a state space equation; according to inversion sliding mode control, designing a three-level nesting subsystem to obtain a control input quantity expression; acquiring a control threshold zone of a control input according to the physical characteristics of an aircraft executing mechanism; based on the control input quantity expression, calculating the control input quantity of each moment according to the flight parameters of the aircraft and the target at each moment; identifying whether the ideal input quantity at each moment exceeds a control threshold value band, if so, calculating the external compensation control quantity at each moment, and superposing the external compensation control quantity to the control input quantity to obtain a total control quantity, otherwise, obtaining the control input quantity as the total control quantity; and controlling the flight attitude of the aircraft according to the total control quantity at each moment. The method realizes the rapid approach to the ideal value and leads the guidance control to be more accurate.
Description
Technical Field
The application relates to the technical field of guidance control, in particular to an aircraft guidance control integrated method and system based on sliding mode control.
Background
The hypersonic aircraft in the near space is mainly divided into a booster gliding hypersonic aircraft and a hypersonic cruise missile. Because of the high flying speed and good maneuverability of hypersonic aircrafts in the near space, the hypersonic aircrafts become a focus subject of attention of military countries in the world.
The hypersonic speed aircraft has the characteristics of high flying speed, high cruising height, strong sudden prevention capability, high detection difficulty and the like, and is a difficult problem in the interception field of the existing precise guidance aircraft.
The flight interception task is game of an intercepted task aircraft and an intercepted person, on one hand, the target can make an evasion action to break through interception, and on the other hand, according to the break-through interception maneuver of the target, the aircraft is required to change the self flight state and flight track, and large maneuver overload is required. The traditional method only depends on the pneumatic rudder, and the interception fault-tolerant rate is extremely low due to the self characteristic limitation of the pneumatic rudder and the environmental factors of small high-altitude dynamic pressure.
The conventional sliding mode control based on the inversion method enables a controlled object to be in an oscillation state in the process of continuously approaching to a sliding mode surface, and due to the slow response of the inversion method, when the sliding mode control quantity cannot enable the aircraft to approach to an ideal state in a short time, an ideal target line-of-sight angle cannot be achieved, and the aircraft cannot be controlled well.
Disclosure of Invention
In view of the above analysis, the embodiment of the application aims to provide an integrated method and system for controlling guidance of an aircraft based on sliding mode control, which are used for solving the problem of slow approach of the existing sliding mode control.
In one aspect, the embodiment of the application provides an integrated method for aircraft guidance control based on sliding mode control, which comprises the following steps:
establishing a guidance control integrated model, and converting the guidance control integrated model into a state space equation;
based on a state space equation, designing a three-level nesting subsystem according to inversion sliding mode control to obtain a control input quantity expression;
acquiring a control threshold zone of a control input according to the physical characteristics of an aircraft executing mechanism;
based on the control input quantity expression, calculating the control input quantity of each moment according to the acquired flight parameters of the aircraft and the target at each moment;
identifying whether the ideal input quantity at each moment exceeds a control threshold value band, if so, calculating the external compensation control quantity at each moment, and superposing the external compensation control quantity to the control input quantity to obtain the total control quantity; otherwise, the control input quantity is the total control quantity;
and controlling the flight attitude of the aircraft according to the total control quantity at each moment.
Based on a further improvement of the method, obtaining a control threshold zone of control input according to a physical characteristic of an aircraft actuator comprises:
obtaining a maximum value and a minimum value of the control input quantity according to the physical characteristics of the aircraft actuating mechanism, and obtaining a maximum control threshold value and a minimum control threshold value according to a preset fluctuation value;
and subtracting the maximum control threshold from the maximum value of the control input quantity, and subtracting the minimum control threshold from the minimum value of the control input quantity to obtain the maximum value and the minimum value of the control threshold band.
Based on the further improvement of the method, the ideal input quantity at each moment is obtained according to the target position at each moment, the ideal sight angle is obtained, and then the ideal input quantity is obtained according to the ideal sight angle by calculation based on the transfer function of the aircraft.
Based on a further improvement of the above method, the external compensation control amount at each time is calculated by the following formula:
，
wherein ,representing the ideal input quantity +.>Represents the maximum value of the control threshold band, +.>Control the minimum value of the threshold band, +.>Represents the maximum value of the control input quantity,/->Representing the minimum value of the control input quantity.
Based on the further improvement of the method, the integrated guidance control model is established, and comprises the following steps: according to flight parameters of the aircraft and the target, a relative motion equation of the aircraft and the target in a longitudinal plane is established, and then a longitudinal disturbance equation of the aircraft is combined to obtain a guidance control integrated model expressed by the following formula:
，
wherein ,a value representing the relative speed of the aircraft and the target in a direction perpendicular to the line of sight, +.>Representation->Derivative with respect to time, < >>Representing a control input quantity; />，/> and />Respectively representing modeling error values caused by aerodynamic parameter disturbance, target maneuver and direct force; />Representing the relative distance between the aircraft and the target in the longitudinal plane, < >>Representation->Derivative with respect to time; />Representing the aircraft velocity; />Represents the angle of sight +.>Representation->Derivative with respect to time; />Representing the ballistic dip angle of the aircraft; />Representing lift force of aircraft->Representing the angle of attack of the flight>Representing the partial derivative of lift with respect to angle of attack; />Representing the mass of an aircraft->Representing engine thrust; />Represents pitch rate, +.> and />Representing the partial derivatives of the pitch moment coefficient to the pitch rate and attack angle respectively; />Representing the moment of inertia of the z-axis, +.>、/> and />Representing the dynamics of an aircraft, < >>Indicating the gravitational acceleration.
Based on the further improvement of the method, the guidance control integrated model is converted into a state space equation, and the state space equation is expressed as follows:
，
wherein ,、/> and />Representing system uncertainty item->、/> and />Representing system state variables>Representing the control input quantity->、/> and />Representing the system interference factor.
Based on the further improvement of the method, based on a state space equation, according to inversion sliding mode control, a three-stage nested subsystem is designed, comprising:
each expression in the state space equation is designed into a first-stage subsystem in sequence, the third-stage subsystems are progressive layer by layer, a layer of sliding mode surface is arranged in each stage of subsystems, the pseudo control quantity from the third-stage subsystem to the first-stage subsystem is tracked in sequence through the control input quantity, the pseudo control quantity of the first-stage subsystem is converged to 0, and the pseudo control quantity of the second-stage subsystem and the pseudo control quantity of the third-stage subsystem are calculated by adopting a first-stage filter.
Based on a further improvement of the above method, the pseudo control amounts of the second stage subsystem and the third stage subsystem are represented by the following formulas:
，
wherein , and />Pseudo control amounts respectively representing the second level subsystem and the third level subsystem, +.>Pseudo control amount representing the second level subsystem +.>Derivative with respect to time, < >> and />Sliding mode surfaces respectively representing a first-stage subsystem and a second-stage subsystem +.>、/>、/> and />Is positive constant, sgn (·) represents the sign function,> and />Respectively indicate-> and />Is provided.
Based on a further improvement of the above method, the control input amount expression is expressed by the following formula:
，
wherein ,representing the slip form surface of the third level subsystem, < >> and />Is of normal number>Representation->Is (are) negative feedback part,)>Pseudo-control representing a third level subsystemProduction of->Derivative with respect to time.
In another aspect, an embodiment of the present application provides an integrated guidance control system for an aircraft based on sliding mode control, including:
the model construction module is used for establishing a guidance control integrated model and converting the guidance control integrated model into a state space equation;
the sliding mode control module is used for designing a three-level nested subsystem according to inversion sliding mode control based on a state space equation to obtain a control input quantity expression;
the threshold value band construction module is used for acquiring a control threshold value band of the control input quantity according to the physical characteristics of the aircraft actuating mechanism;
the control input quantity acquisition module is used for calculating the control input quantity at each moment according to the acquired flight parameters of the aircraft and the target at each moment based on the control input quantity expression;
the total control quantity acquisition module is used for identifying whether the ideal input quantity at each moment exceeds a control threshold value band, if so, calculating the external compensation control quantity at each moment, and superposing the external compensation control quantity to the control input quantity to obtain the total control quantity; otherwise, the control input quantity is the total control quantity;
and the aircraft control module is used for controlling the flight attitude of the aircraft according to the total control quantity at each moment.
Compared with the prior art, the application has at least one of the following beneficial effects: the vibration amplitude of the controlled object is limited by setting a control threshold value zone, and when the required control quantity exceeds the control threshold value zone in the ideal view angle state at each moment, the external compensation control quantity at each moment is dynamically calculated, so that the oscillation amplitude is reduced, the problem that an ideal value cannot be reached in a short time due to asymptotic stability of an inversion sliding mode is solved, the aircraft is enabled to quickly and linearly obtain a large overload, and the ideal target view angle is reached, thereby completing control optimization.
In the application, the technical schemes can be mutually combined to realize more preferable combination schemes. Additional features and advantages of the application will be set forth in the description which follows, and in part will be obvious from the description, or may be learned by practice of the application. The objectives and other advantages of the application may be realized and attained by the structure particularly pointed out in the written description and drawings.
Drawings
The drawings are only for purposes of illustrating particular embodiments and are not to be construed as limiting the application, like reference numerals being used to designate like parts throughout the drawings;
FIG. 1 is a flowchart of an integrated method for aircraft guidance control based on sliding mode control in embodiment 1 of the present application;
FIG. 2 is a schematic diagram of the relative motion relationship between an aircraft and a target in embodiment 1 of the present application;
fig. 3 is a schematic structural diagram of an integrated guidance control system for an aircraft based on sliding mode control in embodiment 2 of the present application.
Detailed Description
The following detailed description of preferred embodiments of the application is made in connection with the accompanying drawings, which form a part hereof, and together with the description of the embodiments of the application, are used to explain the principles of the application and are not intended to limit the scope of the application.
Example 1
The application discloses an integrated method for aircraft guidance control based on sliding mode control, which is shown in fig. 1 and comprises the following steps:
s11, establishing a guidance control integrated model, and converting the guidance control integrated model into a state space equation.
The method for establishing the guidance control integrated model comprises the following steps:
(1) as shown in fig. 2, M represents an aircraft, T represents a target, and according to the flight parameters of the aircraft and the target, the following equation of relative motion of the aircraft and the target in the longitudinal plane is established:
，
wherein ,representing the relative distance between the aircraft and the target in the longitudinal plane, < >>Representation->Derivative with respect to time; and />Representing the aircraft and target rates, respectively; />Represents the angle of sight +.>Representation->Derivative with respect to time, < >> and />Representing the ballistic tilt angles of the aircraft and the target, respectively.
(2) Under the assumption of small disturbance, an aircraft longitudinal disturbance equation is constructed as follows:
，
wherein ,representing the moment of inertia of the z-axis, +.>Represents pitch rate, +.>Representation->The derivative with respect to time is given by, and />Representing the partial derivatives of the pitch moment coefficient to the pitch rate and attack angle respectively; />Representing the lift of the aircraft and,representing the angle of attack of the flight>Representing the partial derivative of lift with respect to angle of attack; />Representing the mass of an aircraft->Represents engine thrust +.>Representation->Derivative with respect to time, < >>Representation->Derivative with respect to time, < >>Indicating the gravitational acceleration.
(3) And (3) combining the formulas (1) and (2) to obtain a guidance control integrated model shown as follows:
，
wherein ,a value representing the relative speed of the aircraft and the target in a direction perpendicular to the line of sight, +.>Representation->Derivative with respect to time, < >>Representing a control input quantity; />，/> and />Respectively representing modeling error values caused by aerodynamic parameter disturbance, target maneuver and direct force; />、/> and />Representing the kinetic coefficients of the aircraft.
S12, designing a three-level nested subsystem based on a state space equation according to inversion sliding mode control, and obtaining a control input quantity expression.
Based on formula (3), the arrangementSystem state variables、/> and />The guidance control integrated model is converted into a state space equation as follows:
，
wherein ,、/> and />Representing system uncertainty item->、/> and />Representing system interference factors, +.>，，/>And->、/> and />Is a preset normal number for defining the limits of the system interference factor.
Further, based on the state space equation, according to the inversion sliding mode control, a three-stage nested subsystem is designed, comprising:
each expression in the state space equation is designed into a first-stage subsystem in sequence, the third-stage subsystems are progressive layer by layer, a layer of sliding mode surface is arranged in each stage of subsystems, the pseudo control quantity from the third-stage subsystem to the first-stage subsystem is tracked in sequence through the control input quantity, the pseudo control quantity of the first-stage subsystem is converged to 0, and the pseudo control quantity of the second-stage subsystem and the pseudo control quantity of the third-stage subsystem are calculated by adopting a first-stage filter.
Specifically, the pseudo control amount of the first stage subsystem is setThe goal is->Even if the line-of-sight angular rate converges to 0; sliding mode face of first stage subsystem->Will->Time derivative due to->The method comprises the following steps of:
，
setting a false control amount of a second level subsystemIdeal condition +.>，/>By->To inhibit, according to the sliding mode control principle, the following formula is obtained:
，
wherein , and />Is positive constant, sgn (·) represents the sign function,>representation->Is provided.
Sliding mode surface for setting second-stage subsystemPseudo control quantity->Tracking, will->Time derivative is carried out, and the following steps are obtained:
，
setting a false control amount of a third level subsystemIdeal condition +.>，/>By->To inhibit, according to the sliding mode control principle, the following formula is obtained:
，
wherein , and />Is positive constant, sgn (·) represents the sign function,>representation->Is (are) negative feedback part,)>Pseudo control amount representing the second level subsystem +.>Derivative with respect to time.
Sliding mode surface with third-stage subsystemPseudo control quantity->Tracking, will->Time derivative is carried out, and the following steps are obtained:
，
by->To suppress, the expression of the obtained control input amount is as follows:
，
wherein , and />Is of normal number>Representation->Is (are) negative feedback part,)>Pseudo control amount representing third level subsystem +.>Derivative with respect to time.
S13, acquiring a control threshold zone of the control input quantity according to the physical characteristics of the aircraft actuating mechanism.
Specifically, the maximum value of the control input is obtained from the physical characteristics of the aircraft actuatorAnd minimum->Then according to the preset fluctuation value +.>Obtaining the maximum control threshold +.>And a minimum control threshold +.>；
Subtracting the maximum control threshold from the maximum value of the control input quantity, subtracting the minimum control threshold from the minimum value of the control input quantity, and obtaining a control threshold band by the following formulaMaximum value of>And minimum->：
，
The aircraft actuator is illustratively a pneumatic rudder, and the control input is 20 degrees at maximum and-20 degrees at minimum, and the control threshold band is [ -16,16] if the fluctuation value is set to 0.2, depending on its physical characteristics.
S14, calculating the control input quantity at each moment according to the flight parameters of the aircraft and the target acquired at each moment based on the control input quantity expression.
In particular, during flight, aircraft flight parameters, including the horizontal flight speed of the aircraft, are obtained by aircraft sensors at each momentAnd longitudinal flight speed>Calculating the horizontal acceleration of the aircraft according to the following formulaDegree->Longitudinal acceleration->And ballistic dip->：
，
wherein ,representing the speed of the aircraft, +.> and />Represents the horizontal displacement and the longitudinal displacement, respectively, < >> and />Respectively indicate-> and />Derivative with respect to time.
Real-time detection of target flight parameters, i.e. real-time motion parameters, including horizontal flight speed of the target, at each moment by an aircraft detection radarAnd longitudinal flight speed>The water of the target is calculated according to the following formulaFlat acceleration->Longitudinal acceleration->And ballistic dip->：
，
wherein ,indicating the target rate +.> and />Represents the horizontal displacement and the longitudinal displacement of the target, respectively,> and />Respectively indicate-> and />Derivative with respect to time.
Based on the flight parameters of the aircraft and the target, the relative distance between the aircraft and the target is calculated according to the following formulaRelative speed->And line of sight angle->：
，
The dynamic coefficients of the aircraft based on the results calculated by formulas (12) - (14)、/> and />And aircraft-dependent constant values, calculating +_in the state space equation>、/> and />Then, based on the formulas (5) - (10), the control input amount +/for each moment is calculated>。
S15, identifying whether the ideal input quantity at each moment exceeds a control threshold value band, if so, calculating the external compensation control quantity at each moment, and superposing the external compensation control quantity to the control input quantity to obtain the total control quantity; otherwise, the control input quantity is the total control quantity.
It should be noted that, when the sliding mode is controlled above the sliding mode surface, the controlled object is continuously reduced to be close to the sliding mode surface. Similarly, when the sliding mode is arranged below the sliding mode surface, the sliding mode is increased to be close to the sliding mode surface, so that the whole motion state is discontinuous, and the sliding mode control enables the controlled object to be continuously close to the sliding mode surface and to be in an oscillating state. The inversion method is slow in response and not rapid, so that the inversion synovial membrane control approach speed is slow. If the controlled object approaches to the sliding mode surface for a long time in an oscillating way, and the target possibly makes an evading action, the aircraft is difficult to complete the interception task, so that the embodiment reduces the oscillation amplitude, improves the flexibility and the quick response of the aircraft and completes the control optimization by constructing a dynamic external compensation control quantity.
Specifically, according to the target positions at all times, an ideal sight angle is obtained, and then based on an aircraft transfer function, an ideal input quantity is obtained through calculation according to the ideal sight angle; and (3) identifying whether the ideal input quantity at each moment exceeds a control threshold value zone of the formula (11), if so, calculating the external compensation control quantity at each moment, and superposing the external compensation control quantity to the control input quantity to obtain the total control quantity, otherwise, obtaining the control input quantity as the total control quantity.
When the ideal input quantity isWhen the control threshold value band is exceeded, the external compensation control amount at each time is calculated by the following formula:
，
the total control amount u is calculated by the following formula:
。
s16, controlling the flight attitude of the aircraft according to the total control quantity at each moment.
Specifically, during flight, the actuator of the aircraft is adjusted according to the total control amount at each time, and the flight attitude of the aircraft is controlled.
Compared with the prior art, the aircraft guidance control integrated method based on the sliding mode control builds a guidance control integrated model, and decomposes the system into three subsystems in cascade by using an inversion method, wherein each subsystem is stabilized by setting a pseudo control quantity in combination with a sliding mode control theory, and the control quantity for enabling the total system to be asymptotically stabilized is obtained by back-pushing step by step; the vibration amplitude of the controlled object is limited by setting a control threshold value zone, and when the required control quantity exceeds the control threshold value zone in the ideal view angle state at each moment, the external compensation control quantity at each moment is dynamically calculated, so that the oscillation amplitude is reduced, the problem that an ideal value cannot be reached in a short time due to asymptotic stability of an inversion sliding mode is solved, the aircraft is enabled to quickly and linearly obtain a large overload, and the ideal target view angle is reached, thereby completing control optimization.
Example 2
In another embodiment of the application, an integrated system for aircraft guidance control based on sliding mode control is disclosed, so as to implement the integrated method for aircraft guidance control based on sliding mode control in embodiment 1. The specific implementation of each module is described with reference to the corresponding description in embodiment 1. As shown in fig. 3, the system includes:
the model construction module 101 is used for establishing a guidance control integrated model and converting the guidance control integrated model into a state space equation;
the sliding mode control module 102 is used for designing a three-level nested subsystem according to inversion sliding mode control based on a state space equation to obtain a control input quantity expression;
the threshold value band construction module 103 is used for acquiring a control threshold value band of the control input quantity according to the physical characteristics of the aircraft actuating mechanism;
a control input quantity acquisition module 104, configured to calculate a control input quantity at each moment according to the acquired flight parameters of the aircraft and the target at each moment based on the control input quantity expression;
the total control amount obtaining module 105 is configured to identify whether the ideal input amount at each moment exceeds a control threshold band, and if so, calculate an external compensation control amount at each moment, and superimpose the external compensation control amount to the control input amount to obtain a total control amount; otherwise, the control input quantity is the total control quantity;
the aircraft control module 106 is configured to control a flight attitude of the aircraft according to the total control amount at each moment.
Because the relevant parts of the present embodiment and the foregoing integrated method for controlling the guidance of the aircraft based on sliding mode control can be referred to each other, repeated descriptions are repeated here, and therefore, the description is omitted here. The principle of the system embodiment is the same as that of the method embodiment, so the system embodiment also has the corresponding technical effects of the method embodiment.
Those skilled in the art will appreciate that all or part of the flow of the methods of the embodiments described above may be accomplished by way of a computer program to instruct associated hardware, where the program may be stored on a computer readable storage medium. Wherein the computer readable storage medium is a magnetic disk, an optical disk, a read-only memory or a random access memory, etc.
The present application is not limited to the above-mentioned embodiments, and any changes or substitutions that can be easily understood by those skilled in the art within the technical scope of the present application are intended to be included in the scope of the present application.
Claims (7)
1. The integrated aircraft guidance control method based on sliding mode control is characterized by comprising the following steps of:
establishing a guidance control integrated model, and converting the guidance control integrated model into a state space equation;
based on a state space equation, designing a three-level nesting subsystem according to inversion sliding mode control to obtain a control input quantity expression;
acquiring a control threshold zone of a control input according to the physical characteristics of an aircraft executing mechanism;
based on the control input quantity expression, calculating the control input quantity of each moment according to the acquired flight parameters of the aircraft and the target at each moment;
identifying whether the ideal input quantity at each moment exceeds a control threshold value band, if so, calculating the external compensation control quantity at each moment, and superposing the external compensation control quantity to the control input quantity to obtain the total control quantity; otherwise, the control input quantity is the total control quantity;
controlling the flight attitude of the aircraft according to the total control quantity at each moment;
the establishing the guidance control integrated model comprises the following steps: according to flight parameters of the aircraft and the target, a relative motion equation of the aircraft and the target in a longitudinal plane is established, and then a longitudinal disturbance equation of the aircraft is combined to obtain a guidance control integrated model expressed by the following formula:
，
wherein ,a value representing the relative speed of the aircraft and the target in a direction perpendicular to the line of sight, +.>Representation->Derivative with respect to time, < >>Representing a control input quantity; />，/> and />Respectively representing modeling error values caused by aerodynamic parameter disturbance, target maneuver and direct force; />Representing the relative distance between the aircraft and the target in the longitudinal plane, < >>Representation->Derivative with respect to time; />Representing the aircraft velocity; />Represents the angle of sight +.>Representation->Derivative with respect to time; />Representing the ballistic dip angle of the aircraft; />Representing lift force of aircraft->Representing the angle of attack of the flight>Representing the partial derivative of lift with respect to angle of attack;representing the mass of an aircraft->Representing engine thrust; />Represents pitch rate, +.> and />Representing the partial derivatives of the pitch moment coefficient to the pitch rate and attack angle respectively; />Representing the moment of inertia of the z-axis, +.>、/> and />Representing the dynamics of an aircraft, < >>Representing gravitational acceleration;
the method for obtaining the control threshold zone of the control input quantity according to the physical characteristics of the aircraft actuating mechanism comprises the following steps: obtaining a maximum value and a minimum value of the control input quantity according to the physical characteristics of the aircraft actuating mechanism, and obtaining a maximum control threshold value and a minimum control threshold value according to a preset fluctuation value; subtracting the maximum control threshold from the maximum value of the control input quantity, and subtracting the minimum control threshold from the minimum value of the control input quantity to obtain the maximum value and the minimum value of the control threshold band;
the external compensation control amount at each time is calculated by the following equation:
，
wherein ,representing the ideal input quantity +.>Represents the maximum value of the control threshold band, +.>The minimum value of the threshold band is controlled,represents the maximum value of the control input quantity,/->Representing the minimum value of the control input quantity.
2. The integrated method for controlling the guidance of the aircraft based on the sliding mode control according to claim 1, wherein the ideal input quantity at each moment is obtained according to the target position at each moment, the ideal sight angle is obtained, and the ideal input quantity is obtained according to the ideal sight angle by calculation based on the transfer function of the aircraft.
3. The integrated sliding mode control-based aircraft guidance control method of claim 1, wherein the integrated guidance control model is converted into a state space equation, expressed as follows:
，
wherein ,、/> and />Representing system uncertainty item->、/> and />A system state variable is represented and is used to represent,representing the control input quantity->、/> and />Representing the system interference factor.
4. The integrated sliding mode control-based aircraft guidance control method of claim 3, wherein the designing a three-stage nested subsystem based on state space equations from an inverted sliding mode control comprises:
each expression in the state space equation is designed into a first-stage subsystem in sequence, the third-stage subsystems are progressive layer by layer, a layer of sliding mode surface is arranged in each stage of subsystems, the pseudo control quantity from the third-stage subsystem to the first-stage subsystem is tracked in sequence through the control input quantity, the pseudo control quantity of the first-stage subsystem is converged to 0, and the pseudo control quantity of the second-stage subsystem and the pseudo control quantity of the third-stage subsystem are calculated by adopting a first-stage filter.
5. The integrated sliding mode control-based aircraft guidance control method according to claim 4, wherein the pseudo control amounts of the second stage subsystem and the third stage subsystem are represented by the following formulas:
，
wherein , and />Pseudo control amounts respectively representing the second level subsystem and the third level subsystem, +.>Pseudo control amount representing the second level subsystem +.>Derivative with respect to time, < >> and />Sliding mode surfaces respectively representing a first-stage subsystem and a second-stage subsystem +.>、/>、/> and />Is positive constant, sgn (·) represents the sign function,> and />Respectively indicate-> and />Is provided.
6. The integrated sliding mode control-based aircraft guidance control method according to claim 5, wherein the control input amount expression is expressed by the following formula:
，
wherein ,representing the slip form surface of the third level subsystem, < >> and />Is of normal number>Representation->Is provided with a negative feedback portion of the (c),pseudo control amount representing third level subsystem +.>Derivative with respect to time.
7. An aircraft guidance control integration system based on sliding mode control, characterized by comprising:
the model construction module is used for establishing a guidance control integrated model and converting the guidance control integrated model into a state space equation;
the sliding mode control module is used for designing a three-level nested subsystem according to inversion sliding mode control based on a state space equation to obtain a control input quantity expression;
the threshold value band construction module is used for acquiring a control threshold value band of the control input quantity according to the physical characteristics of the aircraft actuating mechanism;
the control input quantity acquisition module is used for calculating the control input quantity at each moment according to the acquired flight parameters of the aircraft and the target at each moment based on the control input quantity expression;
the total control quantity acquisition module is used for identifying whether the ideal input quantity at each moment exceeds a control threshold value band, if so, calculating the external compensation control quantity at each moment, and superposing the external compensation control quantity to the control input quantity to obtain the total control quantity; otherwise, the control input quantity is the total control quantity;
the aircraft control module is used for controlling the flight attitude of the aircraft according to the total control quantity at each moment;
the establishing the guidance control integrated model comprises the following steps: according to flight parameters of the aircraft and the target, a relative motion equation of the aircraft and the target in a longitudinal plane is established, and then a longitudinal disturbance equation of the aircraft is combined to obtain a guidance control integrated model expressed by the following formula:
，
wherein ,a value representing the relative speed of the aircraft and the target in a direction perpendicular to the line of sight, +.>Representation->Derivative with respect to time, < >>Representing a control input quantity; />，/> and />Respectively representing modeling error values caused by aerodynamic parameter disturbance, target maneuver and direct force; />Representing the relative distance between the aircraft and the target in the longitudinal plane, < >>Representation->Derivative with respect to time; />Representing the aircraft velocity; />Represents the angle of sight +.>Representation->Derivative with respect to time; />Representing the ballistic dip angle of the aircraft; />Representing lift force of aircraft->Representing the angle of attack of the flight>Representing the partial derivative of lift with respect to angle of attack; />Representing the mass of an aircraft->Representing engine thrust; />Represents pitch rate, +.> and />Representing the partial derivatives of the pitch moment coefficient to the pitch rate and attack angle respectively; />Representing the moment of inertia of the z-axis, +.>、/> and />Representing the dynamics of an aircraft, < >>Representing gravitational acceleration;
the method for obtaining the control threshold zone of the control input quantity according to the physical characteristics of the aircraft actuating mechanism comprises the following steps: obtaining a maximum value and a minimum value of the control input quantity according to the physical characteristics of the aircraft actuating mechanism, and obtaining a maximum control threshold value and a minimum control threshold value according to a preset fluctuation value; subtracting the maximum control threshold from the maximum value of the control input quantity, and subtracting the minimum control threshold from the minimum value of the control input quantity to obtain the maximum value and the minimum value of the control threshold band;
the external compensation control amount at each time is calculated by the following equation:
，
wherein ,representing the ideal input quantity +.>Represents the maximum value of the control threshold band, +.>The minimum value of the threshold band is controlled,represents the maximum value of the control input quantity,/->Representing the minimum value of the control input quantity.
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