CN111610794A - Large-attack-angle dynamic inverse control method for fighter based on sliding mode disturbance observer - Google Patents
Large-attack-angle dynamic inverse control method for fighter based on sliding mode disturbance observer Download PDFInfo
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Abstract
The invention discloses a dynamic inverse control method for a large attack angle of a fighter based on a sliding mode interference observer, which relates to the technical field of aviation control. And then, the uncertainty of the dynamic inverse design method is compensated by combining a supercoiled sliding mode disturbance observer, a stabilization controller of the disturbed attitude system of the fighter is designed, and the simulation is proved by a Lyapunov method. By proper selection of controller parameters, the error can be bounded stably. The invention ensures the good tracking performance and stability of the flight control system of the fighter under a large attack angle, ensures that dangerous states such as deep stall, tail spin and the like are changed in time, and has good reference significance for the practical application of engineering.
Description
Technical Field
The invention belongs to the technical field of advanced aviation control, and particularly relates to a large attack angle dynamic inverse control method of a fighter based on a sliding mode disturbance observer.
Background
The fighter plane is a military plane for eliminating enemy planes in the air, is a main machine type for military air combat, and occupies an irreplaceable position in both ground and air combat. The fighter plane has the condition that the attack angle exceeds the stall attack angle, the pneumatic control surface operation efficiency is reduced or even fails in air combat, when the fighter plane enters a large attack angle area, the pneumatic and flight characteristics are greatly changed, such as nonlinearity, asymmetry, cross coupling and the like of aerodynamic force, so that the stability and the operability of the fighter plane are sharply changed, a plurality of special flight phenomena occur, such as wing shake, upward pitch, nose lateral deviation, over stall rotation, deep stall, tail rotation and the like, the flight state is dangerous and uncontrollable, and if the flight state cannot be separated as soon as possible, the fighter plane and the pilot plane can be caused with unexpected serious consequences.
Therefore, the quality of the design performance of the flight control system of the fighter directly influences whether the fighter can win the air battle. The traditional flight control method such as PID control is mostly designed based on a linear model, and the PID control is widely applied to various civil and military flight control system designs all the time due to the characteristics of simple structure, strong robustness, easy realization and the like. However, the fighter plane model is a multivariable system with strong nonlinearity and strong coupling, and is influenced by flight environment and unsteady aerodynamic force when the fighter plane model is subjected to super maneuver, so that strong external interference exists in the system. Thus, the design of control systems for a fighter aircraft passing a stall maneuver faces design challenges not encountered in conventional flight. Through foreign advanced fighter plane experiments and flight tests, the dynamic inverse control method is considered to be an effective method for controlling the fighter plane. Dynamic inverse control requires that the system model be accurately established and the aircraft state must be accurately measurable or estimable. Considering the situations that the model has uncertain modeling, external interference and the like, particularly when the airplane is in super maneuver flight, the aerodynamic parameters are changed violently due to the violent changes of aerodynamic force and aerodynamic moment. Therefore, when the nonlinear dynamic inverse is adopted to control the airplane, the robust characteristic of the airplane is seriously influenced by the dynamic inverse error.
In summary, the prior art lacks an effective control method for the fighter at a large attack angle, cannot meet the rapidity and stability of the fighter flying at the large attack angle, and cannot ensure that the fighter is transformed into dangerous states such as deep stall, tail spin and the like. The dynamic model is subjected to multi-loop layering based on a time scale separation principle, uncertainty of a dynamic inverse design method is compensated by combining a supercoiled sliding mode disturbance observer, a stable controller of a disturbed attitude system of the fighter is designed, and the error can be stably bounded by reasonably selecting parameters of the controller as proved by a Lyapunov method. The simulation result proves that the control method is effective to the control of the disturbed posture system.
Disclosure of Invention
In order to solve the problems in the prior art, the invention provides a dynamic inverse control method for a large attack angle of a fighter based on a sliding mode disturbance observer, which aims at realizing the effect of effective and rapid flight control on the fighter when the fighter carries out super maneuvering action and enters a large attack angle state.
In order to achieve the purpose, the invention adopts the technical scheme that:
a large attack angle dynamic inverse control method of a fighter based on a sliding mode disturbance observer comprises the following steps:
s1, when the fighter is in a large attack angle flight state, according to the principle of time scale separation, the nonlinear state variable of the fighter is decomposed into two groups of variables based on different time scales by taking airflow angle control and attitude angle rate control as requirements, namely the airflow angle variable and the angle rate variable;
s2, establishing an airflow angle loop model, namely a fast loop model, according to the airflow angle variable, solving an airflow angle loop control law based on a dynamic inverse control method, and compensating uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain a stable airflow angle loop controller;
s3, establishing an attitude angular rate loop model, namely a slow loop model, according to the angular rate variable, solving an attitude angular rate loop control law based on a dynamic inverse control method, and compensating uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain a stable angular rate loop controller.
Further, the supercoiled sliding-mode disturbance observer has the following assumptions and lemmas:
assume 3.1: the partial derivative of the fighter model complex disturbance D with respect to time t is continuous and bounded, and there is a known bounded constant Z >0 that holds the following:
assume 3.2: the system state is considerable, the output and reference signals are continuously differentiable and bounded with respect to time;
3.1 of theory: given the disturbed nonlinear differential equation:
wherein ξ (t) is unknown bounded interference, and
ξ (t), C is the upper bound of the interference derivative, x (t) is the state at time t,
is the derivative of x (t), τ is the time constant, x (τ) is the state to be integrated, w1And w2Is a constant coefficient, if
w21.1C or more, then x (t) and its derivatives
Convergence to zero in a finite time, tr≤(7.6x(0))/(w2-C), x (0) being an initial state;
for a multi-input multi-output affine nonlinear uncertain system of a fighter, designing a supercoiled sliding-mode disturbance observer as follows:
wherein s is an auxiliary sliding mode vector, z is an interference observation state quantity,
to disturb the derivative of the observed state quantity, u is the control quantity,
f and g are functions of the state x.
Further, the specific process of step S2 is as follows:
the airflow angle loop model is as follows:
wherein the content of the first and second substances,
respectively as the derivative commands of the attack angle, the sideslip angle and the roll angle around the speed axis,
respectively a roll angle rate derivative instruction, a pitch angle rate derivative instruction and a yaw angle rate derivative instruction,e、a、rrespectively an elevator deflection angle, an aileron deflection angle, a rudder deflection angle, fsRepresenting the amount of resultant force acting on the aircraft that is independent of the steering control plane and angular rate,
is the amount of resultant force acting on the aircraft that is related to angular rate,
the quantity related to the control surface in the resultant force acting on the aircraft is negligible since the resultant force is mainly related to the angular rate;
a slow loop control law is obtained by a dynamic inverse control method based on a sliding mode disturbance observer, and a given airflow angle derivative instruction is used
Superscript T denotes a transpose, inverse-push-out angular rate derivative instruction for an aircraft
The control law expression is:
wherein s is1、z1The slow loop auxiliary sliding mode vector and the air flow angle interference observed quantity, x, are respectively1＝[a β μ]TIn the state of an air flow angle, x2＝[p q r]TIn the angular rate state, a, β and mu are respectively an attack angle, a sideslip angle and a roll angle around a speed axis, p, q and r are respectively a roll angle rate, a pitch angle rate and a yaw angle rate,
is x1cThe derivative of (a) of (b),
is z1Derivative of (a), x2norIs the dynamic inverse component of angular velocity, x2oFor the angular velocity sliding mode disturbance compensation component,
is a slow loop sliding mode control quantity.
Further, the amount f of the resultant force acting on the aircraft that is independent of the steering control plane and angular ratesThe expression of (a) is:
fs＝[fafβfμ]T
wherein f isaTo remove resultant forces affecting the angle of attack in addition to steering control surface and angular rate, fβTo remove resultant forces affecting the sideslip angle in addition to steering control surface and angular rate, fμIn order to remove the resultant external force which influences the track rolling angle besides the control surface and the angular rate.
Further, the amount of the resultant force acting on the aircraft is related to the angular rate
The expression of (a) is:
wherein, gapThe resultant external force, g, that the roll rate has an influence on the angle of attackaqFor resultant external forces of pitch rate affecting angle of attack, garThe resultant external force, g, affecting the angle of attack for yaw rateβpThe resultant external force, g, affecting the sideslip angle for the roll rateβqResultant external force, g, affecting the sideslip angle for the pitch angle rateβrResultant external force, g, affecting sideslip angle for yaw rateμpFor resultant external forces, g, affecting pitch rate on yaw angleμrThe resultant force, g, affecting yaw for yaw rateμqThe resultant external force that the pitch angle rate has an influence on the yaw angle.
Further, the specific process of step S3 is as follows:
the attitude angular rate loop model is:
wherein f isfTo remove part of the attitude angular rate of control surface steering, gfManipulating a derivative portion for the control surface;
fast obtaining method of dynamic inverse control based on sliding mode disturbance observerLoop control law, given angular velocity command x2cThrust against three rudder offsets u of the aircraftc＝[e a r]TThe fast loop control law expression is:
wherein s is2、z2Respectively a fast loop auxiliary sliding mode vector and an attitude angular velocity disturbance observed quantity,
is x2cDerivative of (a), z2In order to interfere with the observation of the state quantity of the airflow angle,
is z2Derivative of unorFor dynamic inverse control, uoIn order to perform the sliding-mode disturbance compensation control,
the control quantity is a fast loop sliding mode control quantity.
Further, the part f of the attitude angular rate of removing the control surface manipulationfThe expression of (a) is:
wherein h isEIs the angular momentum of the engine and is,
Ix、Iy、Izare the moments of inertia about three body axes, IxzIs the product of inertia, m0、n0、l0Respectively, pitch, yaw and roll moments, which remove control plane manipulation.
Further, the rudderSurface manipulation derivative portion gfThe expression of (a) is:
wherein, gpeA derivative of the roll rate for elevator induced; gpaA derivative of roll rate induced for the aileron; gprA derivative of the rudder induced roll rate; gqeThe pitch rate derivative for elevator induced; greA derivative of yaw rate for elevator induced; graA derivative of yaw rate induced for the aileron; grrThe rudder induced yaw rate derivative.
Compared with the prior art, the invention has the following beneficial effects:
the invention utilizes a large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer, does not depend on a balance point design control law in a certain range, and utilizes the nonlinear inverse of the system to cancel the nonlinearity of the controlled system, thereby realizing the global feedback linearization in the large attack angle stall maneuver flight range and successfully realizing the decoupling control. And the disturbance observer is introduced to compensate the inverse error, so that the dependence of the dynamic inverse on the model accuracy can be reduced, the robustness of the system is improved, and the maneuverability and the stability of the fighter under a large attack angle are ensured.
Drawings
FIG. 1 is a schematic diagram of the structural principle of the present invention;
FIG. 2 is a time domain response diagram of the angle of attack in embodiment 1 of the present invention;
FIG. 3 is a sideslip angle time domain response diagram in embodiment 1 of the present invention;
FIG. 4 is a time-domain response diagram of the track roll angle in embodiment 1 of the present invention.
Detailed Description
The present invention will be further described with reference to the following examples.
The invention provides a large attack angle dynamic inverse control method of a fighter based on a sliding mode disturbance observer, the structure of which is schematically shown in figure 1 and comprises a fast loop dynamic inverse controller and a slow loop dynamic inverse controllerThe method comprises the steps of a road dynamic inverse controller, a fast loop sliding mode disturbance observer and a slow loop sliding mode disturbance observer, wherein the observer is used for observing the dynamic inverse of the road dynamic
Is an adder. The input values of the whole fighter control system comprise an attack angle instruction, a sideslip angle instruction and a roll angle instruction around a speed axis, and the output values comprise flight speed, airflow angle, attitude angle rate, flight height and position information.
Taking a certain type (F16) fighter as an example, when the fighter is in a large attack angle flight state, according to the principle of time scale separation, airflow angle control and attitude angle rate control are taken as requirements to decompose nonlinear state variables of the fighter into two groups of variables based on different time scales, namely airflow angle variables and angular rate variables. And establishing an airflow angle loop model according to the airflow angle variable, solving an airflow angle loop control law based on a dynamic inverse control method, and compensating uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain the stable airflow angle loop controller. And establishing an attitude angular rate loop model according to the angular rate variable, solving an attitude angular rate loop control law based on a dynamic inverse control method, and compensating the uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain a stable angular rate loop controller.
A large attack angle dynamic inverse control method of a fighter based on a sliding mode disturbance observer comprises the following steps:
s1, when the fighter is in a large attack angle flight state, according to the principle of time scale separation, the nonlinear state variable of the fighter is decomposed into two groups of variables based on different time scales by taking airflow angle control and attitude angle rate control as requirements, namely the airflow angle variable and the angle rate variable;
s2, establishing an airflow angle loop model, namely a fast loop model, according to the airflow angle variable, solving an airflow angle loop control law based on a dynamic inverse control method, and compensating uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain a stable airflow angle loop controller;
s3, establishing an attitude angular rate loop model, namely a slow loop model, according to the angular rate variable, solving an attitude angular rate loop control law based on a dynamic inverse control method, and compensating uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain a stable angular rate loop controller.
The supercoiled sliding-mode disturbance observer has the following assumptions and lemmas:
assume 3.1: the partial derivative of the fighter model complex disturbance D with respect to time t is continuous and bounded, and there is a known bounded constant Z >0 that holds the following:
assume 3.2: the system state is considerable, the output and reference signals are continuously differentiable and bounded with respect to time;
3.1 of theory: given the disturbed nonlinear differential equation:
wherein ξ (t) is unknown bounded interference, and
ξ (t), C is the upper bound of the interference derivative, x (t) is the state at time t,
is the derivative of x (t), τ is the time constant, x (τ) is the state to be integrated, w1And w2Is a constant coefficient, if
w21.1C or more, then x (t) and its derivatives
In a limited timeInner convergence to zero point, convergence time tr≤(7.6x(0))/(w2-C), x (0) being an initial state;
for a multi-input multi-output affine nonlinear uncertain system of a fighter, designing a supercoiled sliding-mode disturbance observer as follows:
wherein s is an auxiliary sliding mode vector, z is an interference observation state quantity,
to disturb the derivative of the observed state quantity, u is the control quantity,
f and g are functions of the state x.
Commanded by a given derivative of the airflow angle
The superscript T represents transposition, a slow loop control law is obtained based on a dynamic inverse control method of the sliding mode disturbance observer, and an angular velocity derivative instruction of the airplane is reversely deduced
The airflow angle loop model is as follows:
wherein the content of the first and second substances,
respectively as the derivative commands of the attack angle, the sideslip angle and the roll angle around the speed axis,
respectively as roll angle rate derivative instruction and pitch angle rate derivative instructionA command for the derivative of yaw rate and a yaw rate,e、a、rrespectively an elevator deflection angle, an aileron deflection angle, a rudder deflection angle, fsRepresenting the amount of resultant force acting on the aircraft that is independent of the steering control plane and angular rate,
is the amount of resultant force acting on the aircraft that is related to angular rate,
the quantity related to the control surface in the resultant force acting on the aircraft is negligible since the resultant force is mainly related to the angular rate;
the amount f of the resultant force acting on the aircraft which is independent of the control surface and the angular ratesThe expression of (a) is:
fs＝[fαfβfμ]T
wherein f isαTo remove resultant forces affecting the angle of attack in addition to steering control surface and angular rate, fβTo remove resultant forces affecting the sideslip angle in addition to steering control surface and angular rate, fμIn order to remove the resultant external force which influences the track rolling angle besides the control surface and the angular rate.
The angular rate-related quantity of the resultant force acting on the aircraft
The expression of (a) is:
wherein, gαpThe resultant external force, g, that the roll rate has an influence on the angle of attackαqFor resultant external forces of pitch rate affecting angle of attack, gαrThe resultant external force, g, affecting the angle of attack for yaw rateβpThe resultant external force, g, affecting the sideslip angle for the roll rateβqResultant external force, g, affecting the sideslip angle for the pitch angle rateβrResultant external force, g, affecting sideslip angle for yaw rateμpFor resultant external forces, g, affecting pitch rate on yaw angleμrThe resultant force, g, affecting yaw for yaw rateμqThe resultant external force that the pitch angle rate has an influence on the yaw angle.
Finally, a slow loop control law is obtained based on a dynamic inverse control method of the sliding mode disturbance observer, and a given airflow angle derivative instruction is used
Superscript T denotes a transpose, inverse-push-out angular rate derivative instruction for an aircraft
The slow loop control law expression is:
wherein s is1、z1The slow loop auxiliary sliding mode vector and the air flow angle interference observed quantity, x, are respectively1＝[α β μ]TIn the state of an air flow angle, x2＝[p q r]TIn the state of angular rate, α, β and mu are respectively an attack angle, a sideslip angle and a roll angle around a speed axis, p, q and r are respectively a roll angle rate, a pitch angle rate and a yaw angle rate,
is x1cThe derivative of (a) of (b),
is z1Derivative of (a), x2norIs the dynamic inverse component of angular velocity, x2oFor the angular velocity sliding mode disturbance compensation component,
is a slow loop sliding mode control quantity.
The method is obtained by a given angular speed instruction and a dynamic inverse control method based on a sliding mode disturbance observerTo the slow loop control law, three rudder deflection u of the airplane are reversely pushedc＝[e a r]T；
The attitude angular rate loop model is:
wherein f isfTo remove part of the attitude angular rate of control surface steering, gfManipulating a derivative portion for the control surface;
removing part f of attitude angular rate of control surface steeringfThe expression of (a) is:
wherein h isEIs the angular momentum of the engine and is,
Ix、Iy、Izare the moments of inertia about three body axes, IxzIs the product of inertia, m0、n0、l0Respectively, pitch, yaw and roll moments, which remove control plane manipulation.
Control surface control derivative part gfThe expression of (a) is:
wherein, gpeA derivative of the roll rate for elevator induced; gpaA derivative of roll rate induced for the aileron; gprA derivative of the rudder induced roll rate; gqeThe pitch rate derivative for elevator induced; greA derivative of yaw rate for elevator induced; graA derivative of yaw rate induced for the aileron; grrCaused by ruddersYaw rate derivative.
Finally, a fast loop control law is obtained based on a dynamic inverse control method of the sliding mode disturbance observer, and a given angular speed instruction x is used2cThrust against three rudder offsets u of the aircraftc＝[e a r]TThe fast loop control law expression is:
wherein s is2、z2Respectively a fast loop auxiliary sliding mode vector and an attitude angular velocity disturbance observed quantity,
is x2cDerivative of (a), z2In order to interfere with the observation of the state quantity of the airflow angle,
is z2Derivative of unorFor dynamic inverse control, uoIn order to perform the sliding-mode disturbance compensation control,
the control quantity is a fast loop sliding mode control quantity.
Cobra maneuvers were chosen as example 1. After the aircraft stalls, the lateral stability of the aircraft is reduced, and the aircraft is easy to deviate. Roll angle command mu about the speed axiscAnd sideslip angle command βcAll are zero, angle of attack command acIt is also simpler. The time domain simulation result of the attack angle is shown in fig. 2, the time domain simulation result of the sideslip angle is shown in fig. 3, and the time domain simulation result of the track and roll angle is shown in fig. 4.
From the simulation results, the large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer is adopted, and the fighter can accurately track the attack angle command acSideslip angle command βcRoll angle command [ mu ] around speed axiscTherefore, the dynamic inverse control system has better tracking performance and stability.
The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer does not depend on solving or stability analysis of a nonlinear system, only needs to discuss feedback transformation of the system, and therefore has certain universality. The dynamic inverse control design is simple and convenient, the method can be applied to linear and nonlinear systems, the applicability is wide, the sliding mode disturbance observer ensures the global stability of a closed-loop system, and the tracking performance is good.
Aiming at the flight state with large attack angle, the invention adopts a 'time scale separation' method to decompose the state variable of the airplane into two groups of subsystems based on different time scales, and respectively solves the control law by utilizing a dynamic inverse method. And then, the uncertainty of the dynamic inverse design method is compensated by combining a supercoiled sliding mode disturbance observer, a stabilization controller of the disturbed attitude system of the fighter is designed, and the simulation is proved by a Lyapunov method. By proper selection of controller parameters, the error can be bounded stably. The invention ensures the good tracking performance and stability of the flight control system of the fighter under a large attack angle, ensures that dangerous states such as deep stall, tail spin and the like are changed in time, and has good reference significance for the practical application of engineering.
The above description is only of the preferred embodiments of the present invention, and it should be noted that: it will be apparent to those skilled in the art that various modifications and adaptations can be made without departing from the principles of the invention and these are intended to be within the scope of the invention.
Claims (8)
1. A large attack angle dynamic inverse control method of a fighter based on a sliding mode disturbance observer is characterized by comprising the following steps:
s1, when the fighter is in a large attack angle flight state, according to the principle of time scale separation, the nonlinear state variable of the fighter is decomposed into two groups of variables based on different time scales by taking airflow angle control and attitude angle rate control as requirements, namely the airflow angle variable and the angle rate variable;
s2, establishing an airflow angle loop model, namely a fast loop model, according to the airflow angle variable, solving an airflow angle loop control law based on a dynamic inverse control method, and compensating uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain a stable airflow angle loop controller;
s3, establishing an attitude angular rate loop model, namely a slow loop model, according to the angular rate variable, solving an attitude angular rate loop control law based on a dynamic inverse control method, and compensating uncertainty of the dynamic inverse design method by combining a supercoiled sliding mode disturbance observer to obtain a stable angular rate loop controller.
2. The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer according to claim 1, characterized in that: the supercoiled sliding-mode disturbance observer has the following assumptions and lemmas:
assume 3.1: the partial derivative of the fighter model complex disturbance D with respect to time t is continuous and bounded, and there is a known bounded constant Z >0 that holds the following:
assume 3.2: the system state is considerable, the output and reference signals are continuously differentiable and bounded with respect to time;
3.1 of theory: given the disturbed nonlinear differential equation:
wherein ξ (t) is unknown bounded interference, and
ξ (t), C is the upper bound of the interference derivative, x (t) is the state at time t,
is the derivative of x (t)τ is the time constant, x (τ) is the state to be integrated, w1And w2Is a constant coefficient, if
w21.1C or more, then x (t) and its derivatives
Convergence to zero in a finite time, tr≤(7.6x(0))/(w2-C), x (0) being an initial state;
for a multi-input multi-output affine nonlinear uncertain system of a fighter, designing a supercoiled sliding-mode disturbance observer as follows:
wherein s is an auxiliary sliding mode vector, z is an interference observation state quantity,
to disturb the derivative of the observed state quantity, u is the control quantity,
f and g are functions of the state x.
3. The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer according to claim 1, characterized in that: the specific process of step S2 is as follows:
the airflow angle loop model is as follows:
wherein the content of the first and second substances,
respectively as the derivative commands of the attack angle, the sideslip angle and the roll angle around the speed axis,
respectively a roll angle rate derivative instruction, a pitch angle rate derivative instruction and a yaw angle rate derivative instruction,e、a、rrespectively an elevator deflection angle, an aileron deflection angle, a rudder deflection angle, fsRepresenting the amount of resultant force acting on the aircraft that is independent of the steering control plane and angular rate,
is the amount of resultant force acting on the aircraft that is related to angular rate,
the quantity related to the control surface in the combined external force acting on the airplane;
a slow loop control law is obtained by a dynamic inverse control method based on a sliding mode disturbance observer, and a given airflow angle derivative instruction is used
Superscript T denotes a transpose, inverse-push-out angular rate derivative instruction for an aircraft
The slow loop control law expression is:
wherein s is1、z1The slow loop auxiliary sliding mode vector and the air flow angle interference observed quantity, x, are respectively1＝[a β μ]TIn the state of an air flow angle, x2＝[p q r]TIn the angular rate state, a, β and mu are respectively an attack angle, a sideslip angle and a roll angle around a speed axis, p, q and r are respectively a roll angle rate, a pitch angle rate and a yaw angle rate,
is x1cThe derivative of (a) of (b),
is z1Derivative of (a), x2norIs the dynamic inverse component of angular velocity, x2oFor the angular velocity sliding mode disturbance compensation component,
is a slow loop sliding mode control quantity.
4. The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer according to claim 3, characterized in that: the amount f of the resultant force acting on the aircraft which is independent of the control surface and the angular ratesThe expression of (a) is:
fs＝[fafβfμ]T
wherein f isaTo remove resultant forces affecting the angle of attack in addition to steering control surface and angular rate, fβTo remove resultant forces affecting the sideslip angle in addition to steering control surface and angular rate, fμIn order to remove the resultant external force which influences the track rolling angle besides the control surface and the angular rate.
5. The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer according to claim 3, characterized in that: the angular rate-related quantity of the resultant force acting on the aircraft
The expression of (a) is:
wherein, gapThe resultant external force, g, that the roll rate has an influence on the angle of attackαqIs a pitch angleResultant external force with velocity affecting angle of attack, gαrThe resultant external force, g, affecting the angle of attack for yaw rateβpThe resultant external force, g, affecting the sideslip angle for the roll rateβqResultant external force, g, affecting the sideslip angle for the pitch angle rateβrResultant external force, g, affecting sideslip angle for yaw rateμpFor resultant external forces, g, affecting pitch rate on yaw angleμrThe resultant force, g, affecting yaw for yaw rateμqThe resultant external force that the pitch angle rate has an influence on the yaw angle.
6. The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer according to claim 3, characterized in that: the specific process of step S3 is as follows:
the attitude angular rate loop model is:
wherein f isfTo remove part of the attitude angular rate of control surface steering, gfManipulating a derivative portion for the control surface;
a fast loop control law is obtained by a dynamic inverse control method based on a sliding mode disturbance observer, and a given angular speed instruction x is used2cThrust against three rudder offsets u of the aircraftc＝[e a r]TThe fast loop control law expression is:
wherein s is2、z2Respectively a fast loop auxiliary sliding mode vector and an attitude angular velocity disturbance observed quantity,
is x2cDerivative of (a), z2In order to interfere with the observation of the state quantity of the airflow angle,
is z2Derivative of unorFor dynamic inverse control, uoIn order to perform the sliding-mode disturbance compensation control,
the control quantity is a fast loop sliding mode control quantity.
7. The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer according to claim 6, characterized in that: the part f of the attitude angular rate of removing control surface manipulationfThe expression of (a) is:
wherein h isEIs the angular momentum of the engine and is,
Ix、Iy、Izare the moments of inertia about three body axes, IxzIs the product of inertia, m0、n0、l0Respectively, pitch, yaw and roll moments, which remove control plane manipulation.
8. The large attack angle dynamic inverse control method of the fighter based on the sliding mode disturbance observer according to claim 6, characterized in that: the control surface manipulation derivative part gfThe expression of (a) is:
wherein, gpeA derivative of the roll rate for elevator induced; gpaA derivative of roll rate induced for the aileron; gprA derivative of the rudder induced roll rate; gqeThe pitch rate derivative for elevator induced; greA derivative of yaw rate for elevator induced; graA derivative of yaw rate induced for the aileron; grrThe rudder induced yaw rate derivative.
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