CN114035599A - Aircraft attitude control method and device and electronic equipment - Google Patents
Aircraft attitude control method and device and electronic equipment Download PDFInfo
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Abstract
The invention provides an aircraft attitude control method, an aircraft attitude control device and electronic equipment, wherein a state feedback control law mathematical model is determined based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; determining an attitude controller mathematical model containing a fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model; substituting the attitude controller parameter item into a state feedback control law mathematical model, and determining a pneumatic moment coefficient based on a system control quantity mathematical model obtained in advance and the current state value of the aircraft; calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command. The attitude controller mathematical model determined by the mode contains a fractional order differential term, so that the rudder can provide a larger control quantity at the beginning, the rapidity and the accuracy of an aircraft attitude control system are improved, and the convergence time of the aircraft attitude is shortened.
Description
Technical Field
The invention relates to the technical field of attitude control, in particular to an aircraft attitude control method, an aircraft attitude control device and electronic equipment.
Background
A Flight Control System (FCS) is one of important subsystems of the hypersonic aircraft, is an important guarantee that the hypersonic aircraft can smoothly complete a Flight task, and research on a corresponding Flight attitude Control technology becomes one of a core technology and a key technology in the development process of the hypersonic aircraft. The hypersonic aircraft attitude control system has the following tasks: the command of the aircraft executing mechanism is designed, so that the attitude of the aircraft can quickly and accurately track the command signal given by the guidance system, the stability of the aircraft under the action of various uncertainties and external interference is ensured, and a high-precision attitude control system is necessary for realizing the stability and completing the flight task of the aircraft. In a related technology, a control law of fixed time convergence designed based on a time-varying sliding mode method is applied to attitude control of an aircraft loading section, and a system track is located on a zero sliding mode surface at an initial moment by adding a time-varying item, so that the global robustness of an attitude control system is ensured.
Disclosure of Invention
The invention aims to provide an aircraft attitude control method, an aircraft attitude control device and electronic equipment, so that the utilization efficiency of a rudder is improved, and shorter convergence time is set.
The invention provides an aircraft attitude control method, which comprises the following steps: determining a state feedback control law mathematical model based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; wherein, the state feedback control law mathematical model comprises an attitude controller parameter item; determining an attitude controller mathematical model containing a fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model; substituting the attitude controller mathematical model into an attitude controller parameter item in a state feedback control law mathematical model, and determining a pneumatic moment coefficient based on a pre-acquired system control quantity mathematical model and the current state value of the aircraft; calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command.
Further, the step of determining the state feedback control law mathematical model based on a first attitude control mathematical model expressed in a preset format in the pre-acquired nonlinear mathematical model includes: acquiring a second attitude control mathematical model of the six-degree-of-freedom of the aircraft; wherein the second attitude control mathematical model is part of a non-linear mathematical model; converting the second attitude control mathematical model into a first attitude control mathematical model according to a preset format; the first attitude control mathematical model comprises an attitude angle matrix; calculating a first derivative of an attitude angle matrix in the first attitude control mathematical model to obtain a first derivative result; calculating a second derivative of the attitude angle matrix based on the first derivative result to obtain a second derivative result; and determining a state feedback control law mathematical model based on the second derivation result.
Further, the aircraft attitude control method further comprises the following steps: inputting the state feedback control law mathematical model to a second derivation result to obtain a first linear mathematical model; and determining a target linear mathematical model corresponding to the first attitude control mathematical model based on the first linear mathematical model and the first derivation result.
Further, the step of determining the attitude controller mathematical model containing the fractional order differential item by adopting a sliding mode control mode based on the sliding mode surface control mathematical model acquired in advance comprises the following steps: acquiring a sliding mode surface control mathematical model; the sliding mode surface control mathematical model comprises a fractional order operator; calculating a saturation function mathematical model corresponding to the sliding mode surface control mathematical model; and determining the attitude controller mathematical model containing the fractional differentiation item by adopting a synovial membrane control mode based on the saturation function mathematical model.
Further, the step of substituting the attitude controller mathematical model into the attitude controller parameter item in the state feedback control law mathematical model and determining the aerodynamic moment coefficient based on the pre-obtained system control quantity mathematical model and the current state value of the aircraft comprises the following steps: acquiring a current state value of the aircraft; wherein the current state value comprises at least one of: the angle of attack, the sideslip angle, the roll angle rate, the pitch angle rate, and the yaw rate of the aircraft; acquiring a desired state value given by an aircraft guidance instruction, wherein the desired state value comprises at least one of the following: angle of attack, sideslip angle, roll angle of the aircraft; inputting the error between the current state value and the expected state value into a mathematical model of the attitude controller to obtain a first calculation result; substituting the first calculation result into an attitude controller parameter item in a state feedback control law mathematical model to obtain a second calculation result; acquiring a system control quantity mathematical model containing a pneumatic moment parameter item; and taking the second calculation result as a calculation result of the system control quantity mathematical model, and calculating the aerodynamic moment coefficient corresponding to the aerodynamic moment parameter item in the system control quantity mathematical model.
Further, the first attitude control mathematical model is established based on a preset aircraft model adopting a plane-symmetric structure, and the aircraft model comprises: elevon and rudder; the elevon is used for controlling pitching and rolling motions of the aircraft, and comprises a left elevon and a right elevon; the rudder is used for controlling the yaw movement of the aircraft; calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; the step of controlling the attitude of the aircraft based on the rudder deflection control command includes: calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; wherein the rudder deflection control command includes: a first rudder deflection corresponding to the left elevon, a second rudder deflection corresponding to the right elevon, and a third rudder deflection corresponding to the rudder; and inputting the rudder deflection control command into the nonlinear mathematical model to obtain an output result, and controlling the attitude of the aircraft based on the output result.
Further, the first attitude control mathematical model includes at least one of: kinematic models and kinetic models of aircraft.
The invention provides an aircraft attitude control device, which comprises a first determining module, a second determining module and a state feedback control law mathematical model, wherein the first determining module is used for determining the state feedback control law mathematical model based on a first attitude control mathematical model which is obtained in advance and is expressed according to a preset format in a nonlinear mathematical model; wherein, the state feedback control law mathematical model comprises an attitude controller parameter item; the second determination module is used for determining the attitude controller mathematical model containing the fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model; the third determination module is used for substituting the attitude controller mathematical model into an attitude controller parameter item in the state feedback control law mathematical model and determining the aerodynamic moment coefficient based on the pre-acquired system control quantity mathematical model and the current state value of the aircraft; the control module is used for calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command.
The invention provides an electronic device which comprises a processor and a memory, wherein the memory stores machine executable instructions capable of being executed by the processor, and the processor executes the machine executable instructions to realize the aircraft attitude control method.
The invention provides a machine-readable storage medium storing machine-executable instructions which, when invoked and executed by a processor, cause the processor to implement the aircraft attitude control method described above.
The invention provides an aircraft attitude control method, an aircraft attitude control device and electronic equipment, wherein a state feedback control law mathematical model is determined based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; determining an attitude controller mathematical model containing a fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model; substituting the attitude controller parameter item into a state feedback control law mathematical model, and determining a pneumatic moment coefficient based on a system control quantity mathematical model obtained in advance and the current state value of the aircraft; calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command. The attitude controller mathematical model determined by the mode contains a fractional order differential term, so that the rudder can provide a larger control quantity at the beginning, the rapidity and the accuracy of an aircraft attitude control system are improved, and the convergence time of the aircraft attitude is shortened.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
FIG. 1 is a flow chart of a method for controlling aircraft attitude in accordance with an embodiment of the present invention;
FIG. 2 is a schematic structural diagram of an aircraft model with a plane-symmetric winged conical body configuration according to an embodiment of the present invention;
FIG. 3 is a graph of attitude angle versus time provided by an embodiment of the present invention;
FIG. 4 is a plot of rudder deflection versus time provided by an embodiment of the present invention;
fig. 5 is a schematic structural diagram of an aircraft attitude control device according to an embodiment of the present invention;
FIG. 6 is a flow chart of a controller design based on feedback linearization according to an embodiment of the present invention;
fig. 7 is a schematic structural diagram of an electronic device according to an embodiment of the present invention.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the following embodiments, and it should be understood that the described embodiments are some, but not all, embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
At present, the hypersonic aircraft has complex flying environment, four characteristics of strong nonlinear characteristic, strong coupling characteristic, strong time variation and strong uncertain characteristic, the mathematical model is extremely complex, and the attitude control system of the hypersonic aircraft has the following tasks: the command of the aircraft executing mechanism is designed, so that the attitude of the aircraft can quickly and accurately track the command signal given by the guidance system, the stability of the aircraft under the action of various uncertainties and external interference is ensured, and a high-precision attitude control system is necessary for realizing the stability and completing the flight task of the aircraft. In the prior art, an aircraft controller is designed by using a time-varying sliding mode method for attitude control, the rudder changes slowly at the beginning, the utilization rate of the rudder is low, and therefore the rudder needs a larger control quantity at the middle stage to meet the actual requirement and possibly exceeds the physical limit of the rudder. Based on the above, the embodiment of the invention provides an aircraft attitude control method, an aircraft attitude control device and electronic equipment, and the technology can be applied to applications requiring attitude control of an aircraft.
In order to facilitate understanding of the embodiment, a detailed description is first given to an aircraft attitude control method disclosed in the embodiment of the present invention; as shown in fig. 1, the method comprises the steps of:
step S102, determining a state feedback control law mathematical model based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; the state feedback control law mathematical model comprises an attitude controller parameter item.
The first attitude control mathematical model is in a form of a nonlinear Multiple-Input Multiple-Output (MIMO) affine system, and can be obtained in advance, because the first attitude control mathematical model is a nonlinear mathematical model, and each state variable is seriously coupled and is not beneficial to the design of an attitude controller, the first attitude control mathematical model needs to be decoupled, and an auxiliary control variable (corresponding to the parameter item of the attitude controller) is introduced in the decoupling process to design a state feedback control law mathematical model.
And step S104, determining the attitude controller mathematical model containing the fractional order differential item by adopting a sliding mode control mode based on the pre-acquired sliding mode surface control mathematical model.
A fractional order operator is introduced into the sliding mode surface control mathematical model, the fractional order operator has good memory characteristics and differential characteristics, and after the sliding mode surface control mathematical model is obtained in advance, a sliding mode control mode is adopted to design the attitude controller mathematical model, wherein the attitude controller mathematical model comprises a fractional order differential item.
And S106, substituting the attitude controller mathematical model into an attitude controller parameter item in the state feedback control law mathematical model, and determining the aerodynamic moment coefficient based on the pre-acquired system control quantity mathematical model and the current state value of the aircraft.
In practical implementation, the attitude controller mathematical model containing the fractional order differential item is substituted into the auxiliary control variable (corresponding to the attitude controller parameter item) in the state feedback control law mathematical model designed in step S102 to obtain the system control quantity (corresponding to the state feedback control law), and then the system control quantity is substituted into the system control quantity mathematical model obtained in advance, and the aerodynamic moment coefficient is determined by combining the current state value of the aircraft.
Step S108, calculating a rudder deflection control command based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command.
In practical implementation, an aerodynamic data fitting formula and an aerodynamic data interpolation result are utilized to inversely calculate the aerodynamic moment coefficient to obtain a rudder deflection control instruction, the calculated rudder deflection control instruction is input into a nonlinear mathematical model, the rudder is used as an executing mechanism of the aircraft, and the aircraft can directly control the rudder to reach a position shown by the rudder deflection control instruction, so that the flight attitude of the aircraft is changed.
The aircraft attitude control method comprises the steps of determining a state feedback control law mathematical model based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; determining an attitude controller mathematical model containing a fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model; substituting the attitude controller parameter item into a state feedback control law mathematical model, and determining a pneumatic moment coefficient based on a system control quantity mathematical model obtained in advance and the current state value of the aircraft; calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command. The attitude controller mathematical model determined by the mode contains a fractional order differential term, so that the rudder can provide a larger control quantity at the beginning, the rapidity and the accuracy of an aircraft attitude control system are improved, and the convergence time of the aircraft attitude is shortened.
The embodiment of the invention also provides another aircraft attitude control method, which is realized on the basis of the method of the embodiment; the method comprises the following steps:
step 202, determining a state feedback control law mathematical model based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; the state feedback control law mathematical model comprises an attitude controller parameter item.
Specifically, the state feedback control law mathematical model is determined to be obtained through the following steps one to five based on a first attitude control mathematical model expressed according to a preset format in a pre-obtained nonlinear mathematical model:
acquiring a second attitude control mathematical model of the six-degree-of-freedom of the aircraft; wherein the second attitude control mathematical model is part of a non-linear mathematical model.
The attitude control simplified model (corresponding to the second attitude control mathematical model) of the six-degree-of-freedom hypersonic aircraft is a rigid body motion equation set of the 6-degree-of-freedom aircraft established by considering parameter uncertainty and aggregation disturbance, and belongs to a part of a nonlinear mathematical model, and in actual implementation, the attitude control simplified model of the six-degree-of-freedom hypersonic aircraft is as follows:
in the formulas (1) to (6), α is an attack angle, β is a sideslip angle, μ is a roll angle, p is a roll angle rate, q is a pitch angle rate, r is a yaw angle rate, and IxxIs the moment of inertia of the aircraft in the x-axis direction, IyyIs the moment of inertia of the aircraft in the y-axis direction, IzzMoment of inertia, Δ d, of the aircraft in the z-axis1,Δd2,Δd3,Δd4,Δd5,Δd6Is an aggregate perturbation; mxMoment of the aircraft in the direction of the x-axis, MyMoment of the aircraft in the direction of the y-axis, MzIs the moment of the aircraft in the z-axis direction. The attack angle alpha, the sideslip angle beta and the roll angle mu are three variables in a three-dimensional space, and the three angles have a coupling relation as can be known from formulas (1) to (3); as can be seen from equations (4) to (6), the angular velocities p, q, r corresponding to the three angles also have a coupling relationship with each other. It is therefore necessary to control the state variables in the mathematical model for the second attitude: α, β, μ, p, q, r are decoupled.
Step two, converting the second attitude control mathematical model into a first attitude control mathematical model according to a preset format; the first attitude control mathematical model comprises an attitude angle matrix.
For ease of understanding, equations (1) to (6) above are written as a nonlinear Multiple-Input Multiple-Output (MIMO) affine system form (corresponding to the first attitude control mathematical model above):
in formula (7), X ═ α, β, μ, p, q, r]TFor system state variables, moment M ═ Mx,My,Mz]TFor system control quantity, [ Δ d ═ Δ d1,Δd2,Δd3,Δd4,Δd5,Δd6]To aggregate the perturbation vectors, Ω ═ α, β, μ]TFor the attitude angle matrix, f (X) and g (X) are expressed as follows:
the first attitude control mathematical model is established based on a preset aircraft model adopting a plane symmetric structure, and the aircraft model comprises: elevon and rudder; the elevon is used for controlling pitching and rolling motions of the aircraft, and comprises a left elevon and a right elevon; the rudder is used to steer the yaw motion of the aircraft.
In particular, referring to fig. 2, a schematic structural view of a model of an aircraft with a symmetric pyramid with wing (wined-Cone) configuration is shown, which includes a top view and a side view of the aircraft, wherein the top view shows that the aircraft is a planar symmetric structure, the fuselage is an elongated Cone, the wings are triangular, and the model has a pair of horizontal canards 22 and a pair of elevon 24, and the top view shows that the aircraft further includes a rudder 26. The horizontal duck wings 22 are used for improving the longitudinal stability and maneuverability at the subsonic speed, and are folded into the aircraft body at the hypersonic speed stage; the elevon 24 is used to steer the pitch and roll motions of the aircraft, with the left elevon rudder deflection, as seen from the rear to the front, noted as deltaeThe offset of the auxiliary wing rudder of the right elevator is deltaa(ii) a The rudder is used for steering the yaw movement of the aircraft, and the rudder deflection is recorded as deltar。
The main parameters of the aircraft are shown in table 1:
TABLE 1 aircraft Primary parameters
The first attitude control mathematical model includes at least one of: kinematic models and kinetic models of the above-mentioned aircraft.
In the second attitude control mathematical model in the first step, equations (1) to (3) are kinematic models of the aircraft, equations (4) to (6) are kinetic models of the aircraft, and the first attitude control mathematical model is converted from the second attitude control mathematical model according to a preset format, so that the first attitude control mathematical model at least comprises the kinematic models and the kinetic models of the aircraft.
And step three, calculating a first derivative of an attitude angle matrix in the first attitude control mathematical model to obtain a first derivative result.
Since the first attitude control mathematical model is a nonlinear mathematical model, the coupling of the state variables is serious, and the design of the attitude controller is not facilitated, the state variables of the formula (7) (corresponding to the first attitude control mathematical model) are decoupled by a Feedback Linearization (FBL) method. The feedback linearization method is a design method commonly used in the nonlinear system, and the principle thereof is to perform feedback transformation by using the state/output of the system, thereby converting the nonlinear system into a linear system.
In practical implementation, a first derivative is obtained for Ω in equation (7), which can be obtained from equations (1) - (3):
and step four, calculating a second derivative of the attitude angle matrix based on the first derivative result to obtain a second derivative result.
In practical implementation, the second derivative is obtained from Ω in equation (7), which can be obtained from equation (10):
wherein Δ v ═ Δ v1,Δv2,Δv3]TFor uncertainty, the expressions for the remaining functions are as follows:
and step five, determining a state feedback control law mathematical model based on the second derivation result.
The state feedback control law mathematical model is as follows:
M＝E-1(-F+v) (14)
where v is the introduced auxiliary control quantity (corresponding to the above-mentioned attitude controller parameter item), and v ═ v1,v2,v3]E and F correspond to the above equations (13) and (12), respectively.
Inputting the state feedback control law mathematical model to a second-order derivation result to obtain a first linear mathematical model;
the state feedback control law mathematical model, that is, the formula (14) is substituted into the second derivative, that is, the formula (11), and the first linear mathematical model (corresponding to the following formula (15)) is the sum of a linear part v and a nonlinear disturbance part Δ v:
and determining a target linear mathematical model corresponding to the first attitude control mathematical model based on the first linear mathematical model and the first derivation result.
In actual implementation, based on the first linear mathematical model (corresponding to equation (15)) and the first derivative result (corresponding to equation (10)), a target linear mathematical model (corresponding to equation (16)) corresponding to the first attitude control mathematical model is determined, and the target linear mathematical model may be equivalent to a second-order system as follows:
the state variable of this system is Ω, and the control variable is v, Δ v ═ Δ v1,Δv2,Δv3]TIs an uncertainty of the system.
And 204, determining the attitude controller mathematical model containing the fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model.
Specifically, based on a pre-acquired sliding mode surface control mathematical model, determining an attitude controller mathematical model containing a fractional order differential term in a sliding mode by the following steps six to eight:
step six, acquiring a sliding mode surface control mathematical model; the sliding mode surface control mathematical model comprises a fractional order operator.
In practical implementation, the controller sliding mode surface of the hypersonic aircraft (corresponding to the sliding mode surface control mathematical model) is as follows:
wherein, tfFor arbitrarily set convergence time, t0The time of the initial moment of the aircraft is generally 0, t represents the current time value of the aircraft in the motion process, n and q are sliding mode surface gains and are any given values, and DλIs a fractional operator, λ is the order of the fractional order, D is a sign,
presentation pair
The derivative of the order of lambda is calculated,
for tracking errors, where Ω is the actual value of the attitude angle, ΩcFor the desired value of the attitude angle,
is the derivative of the tracking error, K1＝diag[k11,k12,k13]Is a fractional order term gain, K2Determined by the initial value of the system, K2Comprises the following steps:
and step seven, calculating a saturation function mathematical model corresponding to the sliding mode surface control mathematical model.
In order to solve the buffeting problem of the sliding mode control, a saturation function sat (S) is adopted as a switching function of the controller. The saturation function sat(s) is shown below.
Where ε represents the thickness of the boundary layer and is a predetermined value, and S corresponds to the above formula (17).
And step eight, determining the attitude controller mathematical model containing the fractional order differential term by adopting a slip film control mode based on the saturation function mathematical model.
By using a sliding mode control method, the obtained attitude controller (corresponding to the attitude controller mathematical model containing the fractional order differential term) of the hypersonic aircraft is as follows:
wherein K3＝diag[k31,k32,k33]Is a switching gain and has k3i≥Δvi,i＝1,2,3。
In order to facilitate understanding of the above embodiments, initial simulation values and simulation parameters of the hypersonic aircraft are given as shown in table 2:
TABLE 2 hypersonic flight vehicle simulation initial values and simulation parameters
In Table 2, the initial values of the three attitude angles are [0,0 ]]Tdeg, expected value of [3,2,1 ]]Tdeg. When the convergence time t is setfFor 1s,1.2s and 1.5s, respectively, the controller with fractional order shown in formula (19) and the conventional controller without fractional order and with fixed time convergence are used for simulation, and the obtained change of the attitude angle with time is shown in fig. 3, and the change of the rudder deflection with time is shown in fig. 4, wherein the subscript of "1" represents the simulation result of the conventional controller without fractional order, and the subscript of "2" represents the simulation result of the controller with fractional order designed by formula (19).
As shown in fig. 3 and 4, under the same initial conditions, the controller with fractional order designed by formula (19) can provide a larger control amount at the beginning to improve the utilization efficiency of the rudder according to the actual requirement due to the action of the fractional order differential term, while when the fractional order term is not added, the rudder changes slowly at the beginning, the utilization rate of the rudder is lower, which requires the rudder to need a larger control amount at the middle stage to meet the actual requirement, and the control amount may exceed the physical limit of the rudder, such as t in fig. 3 and 4fAt 1.0s, the control requirement is not met if the rudder does not exceed the physical limit, e.g., when tfWhen 1.2s, there is no pointA conventional controller of several orders may also have a relatively slow change in the early stage, resulting in a relatively large change in the rudder at the last moment, which may reduce the control accuracy. Otherwise, the controller with fractional order designed by equation (19) can have a shorter convergence time because of a higher utilization of the rudder.
And step 206, substituting the attitude controller mathematical model into an attitude controller parameter item in the state feedback control law mathematical model, and determining the aerodynamic moment coefficient based on the pre-acquired system control quantity mathematical model and the current state value of the aircraft.
Specifically, the attitude controller mathematical model is substituted into an attitude controller parameter item in the state feedback control law mathematical model, and based on the pre-obtained system control quantity mathematical model and the current state value of the aircraft, the aerodynamic moment coefficient is determined and obtained through the following nine steps to the thirteen steps:
step nine, acquiring a current state value of the aircraft; wherein the current state value comprises at least one of: the angle of attack, the side slip angle, the roll rate, the pitch rate, and the yaw rate of the aircraft.
In practical implementation, the first attitude control mathematical model of the aircraft may obtain, in real time, the value of the state variable of the aircraft at the current time (corresponding to the current state value). The values of the current state variables include at least one of: α, β, μ, p, q, r; where α, β, μ constitute the actual value of the attitude angle.
Step ten, acquiring an expected state value given by an aircraft guidance instruction, wherein the expected state value comprises at least one of the following: angle of attack, sideslip angle, roll angle of the aircraft.
The expected state value is given by guidance instructions of the aircraft, and the expected state value comprises at least one of the following: alpha is alphac,βc,μcThe desired state value αc,βc,μcIs the expected value of attitude angle
And step eleven, inputting the error between the current state value and the expected state value into the attitude controller mathematical model to obtain a first calculation result.
The actual values alpha, beta, mu of the attitude angle are compared with the expected value alphac,βc,μcThe error between the two is substituted as an input term into an attitude controller (corresponding to the above-mentioned attitude controller mathematical model) of the hypersonic flight vehicle shown in formula (19), and the obtained v is the above-mentioned first calculation result.
And step twelve, substituting the first calculation result into the attitude controller parameter item in the state feedback control law mathematical model to obtain a second calculation result.
The attitude controller parameter term is v in a state feedback control law mathematical model shown in formula (14), v (corresponding to the calculation result) obtained in step eleven is substituted into the formula (14), and the control moment M required by the aircraft is calculated and is the second calculation result.
Step thirteen, acquiring a system control quantity mathematical model containing a pneumatic moment parameter item; and taking the second calculation result as a calculation result of the system control quantity mathematical model, and calculating the aerodynamic moment coefficient corresponding to the aerodynamic moment parameter item in the system control quantity mathematical model.
In practical implementation, the moment formula (corresponding to the above system control quantity mathematical model) of the aircraft is:
wherein M ═ Mx,My,Mz]Is the aerodynamic moment, rho is the atmospheric density, S is the reference area of the aircraft, L is the reference length of the aircraft, V is the speed of the aircraft at the current moment, CMIs the aerodynamic moment coefficient of the aircraft (corresponding to the aerodynamic moment parameter term). Substituting the control moment M (corresponding to the second calculation result) required by the aircraft calculated in the step twelve into the formula (21) to reversely calculate the aerodynamic moment coefficient CM。
Step 208, calculating a rudder deflection control command based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command.
Specifically, a rudder deflection control command is calculated based on the aerodynamic moment coefficient; controlling the attitude of the aircraft based on the rudder deflection control command is obtained through the following steps fourteen to fifteen:
step fourteen, calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; wherein the rudder deflection control command includes: the first rudder deflection corresponding to the left elevon, the second rudder deflection corresponding to the right elevon, and the third rudder deflection corresponding to the rudder.
The first rudder is deviated to a left auxiliary elevator wing rudder deviation deltaeThe second rudder is offset delta of the auxiliary wing rudder of the right elevatora(ii) a The third rudder deviates to the rudder deviation deltarUtilizing a pneumatic data fitting formula and a pneumatic data interpolation result to obtain a pneumatic moment coefficient CMInverse calculation of rudder deflection control command deltae，δa，δr。
And step fifteen, inputting the rudder deflection control instruction into the nonlinear mathematical model to obtain an output result, and controlling the attitude of the aircraft based on the output result.
The rudder deflection control command δe，δa，δrInputting a nonlinear mathematical model of the aircraft to enable the attitude angle alpha, beta and mu of the aircraft to accurately track the upper expected attitude angle alphac,βc,μc。
In practical implementation, the nonlinear mathematical model includes, in addition to the second attitude control mathematical model, a centroid motion equation:
wherein equations (22) - (24) are centroid kinematic equations, equations (25) - (27) are centroid kinetic equations, x, y, z are position coordinates in a three-dimensional space, wherein y represents height, the atmospheric density ρ at the current height can be calculated by using a standard atmospheric table, V is velocity, g is gravitational acceleration, θ, ψ are calculatedvAnd gammavRespectively pitch angle, yaw angle and roll angle of speed, FX、FYAnd FZAre the drag, lift and lateral forces of the aircraft.
To facilitate understanding of the above embodiments, referring to a flow chart of a controller design based on feedback linearization shown in fig. 5, first, the nonlinear multiple-input multiple-output control system (corresponding to the first attitude control mathematical model) with severely coupled state variables shown in equation (7) is simplified into a linear control system (corresponding to the target linear mathematical model) shown in equation (16) by using a method of feedback linearization. Then, a posture control law is designed for the linear system shown in the formula (16) by using a sliding mode control method, and the obtained controller (corresponding to a posture controller mathematical model) is shown in a formula (19).
The nonlinear mathematical model of the aircraft can obtain the values of the state variables alpha, beta, mu, p, q, r, rho and V of the aircraft at the current moment in real time, the obtained values of alpha, beta, mu, p, q and r are substituted into a controller (corresponding to an attitude controller mathematical model) shown in a formula (19), an auxiliary control variable V is calculated, the state feedback control law mathematical model shown in a formula (14) is used for solving the control moment M required by the aircraft, and then the values of rho and V obtained by the nonlinear mathematical model are substituted into the moment of the aircraft shown in a formula (21)Calculating aerodynamic moment coefficient C by formula (corresponding to mathematical model of system control quantity)MUtilizing a pneumatic data fitting formula and a pneumatic data interpolation result in the prior art to obtain a pneumatic moment coefficient CMInverse calculation of rudder deflection control command deltae，δa，δrInputting a nonlinear mathematical model of the aircraft, calculating the aerodynamic coefficient and a new moment coefficient at the current moment by using the aerodynamic data fitting formula and the aerodynamic data interpolation result in the prior art again, substituting the aerodynamic coefficient and the aerodynamic moment formula to calculate aerodynamic force and aerodynamic moment, wherein the aerodynamic moment formula is a corresponding formula (21), and the aerodynamic force formula is as follows:
wherein F ═ FX,FY,FZ]Is aerodynamic force, CRIs the aerodynamic coefficient; and (3) substituting the aerodynamic moment into the second attitude control mathematical model, and substituting the aerodynamic force into equations (22) - (27) to obtain the value of the state variable of the aircraft at the current moment so as to control the attitude of the aircraft.
An embodiment of the present invention further provides an aircraft attitude control device, as shown in fig. 6, the device includes: a first determining module 60, configured to determine a state feedback control law mathematical model based on a first attitude control mathematical model expressed in a preset format in a pre-acquired nonlinear mathematical model; wherein, the state feedback control law mathematical model comprises an attitude controller parameter item; the second determining module 61 is configured to determine, based on a sliding mode surface control mathematical model obtained in advance, an attitude controller mathematical model including a fractional order differential item by using a sliding mode control manner; a third determining module 62, configured to substitute the attitude controller mathematical model into an attitude controller parameter item in the state feedback control law mathematical model, and determine an aerodynamic moment coefficient based on a pre-obtained system control quantity mathematical model and a current state value of the aircraft; the control module 63 is used for calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command.
The aircraft attitude control device determines a state feedback control law mathematical model based on a first attitude control mathematical model expressed according to a preset format in a nonlinear mathematical model acquired in advance; determining an attitude controller mathematical model containing a fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model; substituting the attitude controller parameter item into a state feedback control law mathematical model, and determining a pneumatic moment coefficient based on a system control quantity mathematical model obtained in advance and the current state value of the aircraft; calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; and controlling the attitude of the aircraft based on the rudder deflection control command. The attitude controller mathematical model determined by the mode contains a fractional order differential term, so that the rudder can provide a larger control quantity at the beginning, the rapidity and the accuracy of an aircraft attitude control system are improved, and the convergence time of the aircraft attitude is shortened.
Further, the first determining module is configured to: acquiring a second attitude control mathematical model of the six-degree-of-freedom of the aircraft; wherein the second attitude control mathematical model is part of a non-linear mathematical model; converting the second attitude control mathematical model into a first attitude control mathematical model according to a preset format; the first attitude control mathematical model comprises an attitude angle matrix; calculating a first derivative of an attitude angle matrix in the first attitude control mathematical model to obtain a first derivative result; calculating a second derivative of the attitude angle matrix based on the first derivative result to obtain a second derivative result; and determining a state feedback control law mathematical model based on the second derivation result.
Further, the first determining module is further configured to: inputting the state feedback control law mathematical model to a second-order derivation result to obtain a first linear mathematical model; and determining a target linear mathematical model corresponding to the first attitude control mathematical model based on the first linear mathematical model and the first derivation result.
Further, the second determining module is configured to: acquiring a sliding mode surface control mathematical model; the sliding mode surface control mathematical model comprises a fractional order operator; calculating a saturation function mathematical model corresponding to the sliding mode surface control mathematical model; and determining the attitude controller mathematical model containing the fractional differentiation item by adopting a synovial membrane control mode based on the saturation function mathematical model.
Further, the third determining module is further configured to: acquiring a current state value of the aircraft; wherein the current state value comprises at least one of: the angle of attack, the sideslip angle, the roll angle rate, the pitch angle rate, and the yaw rate of the aircraft; acquiring a desired state value given by an aircraft guidance instruction, wherein the desired state value comprises at least one of the following: angle of attack, sideslip angle, roll angle of the aircraft. Inputting the error between the current state value and the expected state value into a mathematical model of the attitude controller to obtain a first calculation result; substituting the first calculation result into an attitude controller parameter item in a state feedback control law mathematical model to obtain a second calculation result; acquiring a system control quantity mathematical model containing a pneumatic moment parameter item; and taking the second calculation result as a calculation result of the system control quantity mathematical model, and calculating the aerodynamic moment coefficient corresponding to the aerodynamic moment parameter item in the system control quantity mathematical model.
Further, the first attitude control mathematical model is established based on a preset aircraft model adopting a plane-symmetric structure, and the aircraft model comprises: elevon and rudder; the elevon is used for controlling pitching and rolling motions of the aircraft, and comprises a left elevon and a right elevon; the rudder is used for controlling the yaw movement of the aircraft; the control module is used for: calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; wherein the rudder deflection control command includes: a first rudder deflection corresponding to the left elevon, a second rudder deflection corresponding to the right elevon, and a third rudder deflection corresponding to the rudder; and inputting the rudder deflection control command into the nonlinear mathematical model to obtain an output result, and controlling the attitude of the aircraft based on the output result.
Further, the first attitude control mathematical model includes at least one of: kinematic models and kinetic models of aircraft.
The implementation principle and the generated technical effects of the aircraft attitude control device provided by the embodiment of the invention are the same as those of the aircraft attitude control method embodiment, and the corresponding contents in the aircraft attitude control method embodiment can be referred to in the aircraft attitude control device embodiment.
An embodiment of the present invention further provides an electronic device, which is shown in fig. 7 and includes a processor 130 and a memory 131, where the memory 131 stores machine executable instructions that can be executed by the processor 130, and the processor 130 executes the machine executable instructions to implement an aircraft attitude control method.
Further, the electronic device shown in fig. 7 further includes a bus 132 and a communication interface 133, and the processor 130, the communication interface 133, and the memory 131 are connected through the bus 132.
The Memory 131 may include a high-speed Random Access Memory (RAM) and may also include a non-volatile Memory (non-volatile Memory), such as at least one disk Memory. The communication connection between the network element of the system and at least one other network element is realized through at least one communication interface 133 (which may be wired or wireless), and the internet, a wide area network, a local network, a metropolitan area network, and the like can be used. The bus 132 may be an ISA bus, PCI bus, EISA bus, or the like. The bus may be divided into an address bus, a data bus, a control bus, etc. For ease of illustration, only one double-headed arrow is shown in FIG. 7, but this does not indicate only one bus or one type of bus.
The processor 130 may be an integrated circuit chip having signal processing capabilities. In implementation, the steps of the above method may be performed by integrated logic circuits of hardware or instructions in the form of software in the processor 130. The Processor 130 may be a general-purpose Processor, and includes a Central Processing Unit (CPU), a Network Processor (NP), and the like; the device can also be a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field Programmable Gate Array (FPGA) or other Programmable logic device, a discrete Gate or transistor logic device, or a discrete hardware component. The various methods, steps and logic blocks disclosed in the embodiments of the present invention may be implemented or performed. A general purpose processor may be a microprocessor or the processor may be any conventional processor or the like. The steps of the method disclosed in connection with the embodiments of the present invention may be directly implemented by a hardware decoding processor, or implemented by a combination of hardware and software modules in the decoding processor. The software module may be located in ram, flash memory, rom, prom, or eprom, registers, etc. storage media as is well known in the art. The storage medium is located in the memory 131, and the processor 130 reads the information in the memory 131 and completes the steps of the method of the foregoing embodiment in combination with the hardware thereof.
The embodiment of the present invention further provides a machine-readable storage medium, where the machine-readable storage medium stores machine-executable instructions, and when the machine-executable instructions are called and executed by a processor, the machine-executable instructions cause the processor to implement the aircraft attitude control method, and specific implementation may refer to method embodiments, and is not described herein again.
The aircraft attitude control method, the aircraft attitude control device and the computer program product of the electronic device provided by the embodiments of the present invention include a computer-readable storage medium storing program codes, instructions included in the program codes may be used to execute the methods described in the foregoing method embodiments, and specific implementations may refer to the method embodiments and are not described herein again.
The functions, if implemented in the form of software functional units and sold or used as a stand-alone product, may be stored in a computer readable storage medium. Based on such understanding, the technical solution of the present invention may be embodied in the form of a software product, which is stored in a storage medium and includes instructions for causing a computer device (which may be a personal computer, a server, or a network device) to execute all or part of the steps of the method according to the embodiments of the present invention. And the aforementioned storage medium includes: a U-disk, a removable hard disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a magnetic disk or an optical disk, and other various media capable of storing program codes.
Finally, it should be noted that: the above embodiments are only used to illustrate the technical solution of the present invention, and not to limit the same; while the invention has been described in detail and with reference to the foregoing embodiments, it will be understood by those skilled in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some or all of the technical features may be equivalently replaced; and the modifications or the substitutions do not make the essence of the corresponding technical solutions depart from the scope of the technical solutions of the embodiments of the present invention.
Claims (10)
1. A method of aircraft attitude control, the method comprising:
determining a state feedback control law mathematical model based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; wherein, the state feedback control law mathematical model comprises an attitude controller parameter item;
determining an attitude controller mathematical model containing a fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model;
substituting the attitude controller mathematical model into an attitude controller parameter item in the state feedback control law mathematical model, and determining a pneumatic moment coefficient based on a pre-obtained system control quantity mathematical model and the current state value of the aircraft;
calculating a rudder deflection control command based on the aerodynamic moment coefficient; controlling the attitude of the aircraft based on the rudder deflection control command.
2. The method according to claim 1, wherein the step of determining the state feedback control law mathematical model based on a first attitude control mathematical model expressed in a preset format among the pre-acquired nonlinear mathematical models comprises:
acquiring a second attitude control mathematical model of the six-degree-of-freedom of the aircraft; wherein the second attitude control mathematical model is part of the non-linear mathematical model;
converting the second attitude control mathematical model into the first attitude control mathematical model according to a preset format; wherein the first attitude control mathematical model comprises an attitude angle matrix;
calculating a first derivative of the attitude angle matrix in the first attitude control mathematical model to obtain a first derivative result;
calculating a second derivative of the attitude angle matrix based on the first derivative result to obtain a second derivative result;
and determining the state feedback control law mathematical model based on the second derivation result.
3. The method of claim 2, further comprising:
inputting the state feedback control law mathematical model to the second derivation result to obtain a first linear mathematical model;
and determining a target linear mathematical model corresponding to the first attitude control mathematical model based on the first linear mathematical model and the first derivation result.
4. The method according to claim 1, wherein the step of determining the attitude controller mathematical model including a fractional derivative term in a slip-mode control manner based on a pre-acquired slip-mode surface control mathematical model comprises:
acquiring the sliding mode surface control mathematical model; the sliding mode surface control mathematical model comprises a fractional order operator;
calculating a saturation function mathematical model corresponding to the sliding mode surface control mathematical model;
and determining the attitude controller mathematical model containing fractional differentiation items by adopting a synovial membrane control mode based on the saturation function mathematical model.
5. The method of claim 1, wherein substituting the attitude controller mathematical model into an attitude controller parameter term in the state feedback control law mathematical model, the step of determining an aerodynamic moment coefficient based on a pre-acquired system control quantity mathematical model and a current state value of the aircraft comprises:
obtaining the current state value of the aircraft; wherein the current state value comprises at least one of: the angle of attack, the sideslip angle, the roll angle rate, the pitch angle rate and the yaw angle rate of the aircraft;
obtaining a desired state value given by the aircraft guidance instruction, wherein the desired state value comprises at least one of: an angle of attack, a sideslip angle, a roll angle of the aircraft;
inputting the error between the current state value and the expected state value into the attitude controller mathematical model to obtain a first calculation result;
substituting the first calculation result into an attitude controller parameter item in the state feedback control law mathematical model to obtain a second calculation result;
acquiring a system control quantity mathematical model containing a pneumatic moment parameter item;
and calculating the aerodynamic moment coefficient corresponding to the aerodynamic moment parameter term in the system control quantity mathematical model by taking the second calculation result as the calculation result of the system control quantity mathematical model.
6. The method of claim 1, wherein the first attitude control mathematical model is established based on a preset aircraft model adopting a plane-symmetric structure, the aircraft model comprising: elevon and rudder; wherein the elevon is used to manipulate pitch and roll motions of the aircraft, the elevon comprising a left elevon and a right elevon; the rudder is used for manipulating the yaw movement of the aircraft;
calculating a rudder deflection control command based on the aerodynamic moment coefficient; the step of controlling the attitude of the aircraft based on the rudder deflection control command includes:
calculating a rudder deflection control command based on the aerodynamic moment coefficient; wherein the rudder deflection control command includes: a first rudder deflection corresponding to the left elevon, a second rudder deflection corresponding to the right elevon, and a third rudder deflection corresponding to the rudder;
and inputting the rudder deflection control command to the nonlinear mathematical model to obtain an output result, and controlling the attitude of the aircraft based on the output result.
7. The method of claim 1, wherein the first attitude control mathematical model includes at least one of: a kinematic model and a kinetic model of the aircraft.
8. An aircraft attitude control device, characterized in that the device comprises:
the first determining module is used for determining a state feedback control law mathematical model based on a first attitude control mathematical model expressed according to a preset format in a pre-acquired nonlinear mathematical model; wherein, the state feedback control law mathematical model comprises an attitude controller parameter item;
the second determination module is used for determining the attitude controller mathematical model containing the fractional order differential item by adopting a sliding mode control mode based on a pre-acquired sliding mode surface control mathematical model;
the third determination module is used for substituting the attitude controller mathematical model into an attitude controller parameter item in the state feedback control law mathematical model and determining a pneumatic moment coefficient based on a pre-acquired system control quantity mathematical model and the current state value of the aircraft;
the control module is used for calculating a rudder deflection control instruction based on the aerodynamic moment coefficient; controlling the attitude of the aircraft based on the rudder deflection control command.
9. An electronic device comprising a processor and a memory, the memory storing machine executable instructions executable by the processor, the processor executing the machine executable instructions to implement the aircraft attitude control method of any one of claims 1 to 7.
10. A machine-readable storage medium having stored thereon machine-executable instructions which, when invoked and executed by a processor, cause the processor to implement the aircraft attitude control method of any one of claims 1 to 7.
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