CN116360258A - Hypersonic deformed aircraft anti-interference control method based on fixed time convergence - Google Patents
Hypersonic deformed aircraft anti-interference control method based on fixed time convergence Download PDFInfo
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Abstract
The invention relates to a hypersonic deformation aircraft anti-interference control method based on fixed time convergence, which comprises the following steps: s1, constructing a gesture motion and pneumatic model based on a geometric model of a hypersonic deformation aircraft; s2, constructing an interference observer for estimating the hypersonic deformed aircraft in real time based on total interference in the instruction flight process based on the attitude motion and the pneumatic model; s3, converting the gesture motion and the pneumatic model into control models facing control, and designing a sliding mode controller converging at fixed time based on real-time estimation results of the disturbance observer, wherein the sliding mode controller tracks instructions of the hypersonic deformation aircraft based on the control models, so that gesture control tasks under disturbance conditions are completed.
Description
Technical Field
The invention relates to the field of aircraft control, in particular to a hypersonic deformed aircraft anti-interference control method based on fixed time convergence.
Background
The deformed aircraft can flexibly change the appearance according to the requirements of flight tasks, so that the flight requirements of a large airspace and a large speed area are met, and the optimal flight is kept, so that the aircraft can complete tasks which cannot be completed by conventional aircraft. The deformed aircraft achieves good combat usability in the area with wide speed and altitude variation range for a new generation of aerospace aircraft flying across the atmosphere. The hypersonic deformed aircraft is a hypersonic aircraft which can actively change the appearance structure according to the requirements of flight environment and flight tasks so as to obtain better aerodynamic characteristics and maneuvering capability, can meet the flight requirements of a large airspace and a large speed domain and is beneficial to improving the flight performance. The hypersonic deformed aircraft takes the appearance parameters as controllable variables, and changes the performance of the aircraft by utilizing the influence of the appearance parameters on the aerodynamic characteristics, so that the hypersonic deformed aircraft can adapt to a flight airspace and a speed domain in a wider range, and further can adapt to more complex flight tasks and flight environments. Meanwhile, the appearance, the flight performance and the like are flexibly changed according to the change of battlefield environment and battlefield tasks, the range, the burst protection performance and the accuracy of the aircraft are enhanced, and the battlefield efficiency and the efficiency-cost ratio of the aircraft can be greatly improved.
The characteristics of fast time variation, nonlinearity, strong coupling, shape variation and the like require the control system to have stronger adaptability, while the uncertainty factors such as unmodeled dynamics, external interference, pneumatic perturbation and the like require the controller to have extremely strong robustness. Conventional adaptive control, robust control and sliding mode control are typically based on a worst case design that may be achieved to achieve some passive immunity, but sacrifice control performance under nominal conditions and control effects can deteriorate rapidly when actual disturbances exceed pre-assumed limits. The controller designed based on the interference observer can actively estimate the interference value and compensate the interference value in the control quantity, and the control performance under the nominal and deviation conditions is considered. The method is characterized in that the method comprises the steps of determining the deformation condition of the hypersonic deformed aircraft, and determining the deformation condition of the hypersonic deformed aircraft according to the deformation condition, wherein the deformation condition is determined by the deformation condition, and the deformation condition is determined by the deformation condition.
Disclosure of Invention
The invention aims to provide an anti-interference control method for a hypersonic deformed aircraft based on fixed time convergence, which aims at the problems of control of reentry sections of hypersonic deformed aircraft and improvement of time convergence performance under interference conditions in the prior art.
In order to achieve the above object, the present invention provides a hypersonic deformed aircraft anti-interference control method based on fixed time convergence, which is characterized by comprising:
s1, constructing a gesture motion and pneumatic model based on a geometric model of a hypersonic deformation aircraft;
s2, constructing an interference observer for estimating the hypersonic deformed aircraft in real time based on total interference in the instruction flight process based on the attitude motion and the pneumatic model;
s3, converting the gesture motion and the pneumatic model into control models facing control, and designing a sliding mode controller converging at fixed time based on real-time estimation results of the disturbance observer, wherein the sliding mode controller tracks instructions of the hypersonic deformation aircraft based on the control models, so that gesture control tasks under disturbance conditions are completed.
According to one aspect of the invention, in step S1, the step of constructing the attitude motion and aerodynamic model based on the geometric model of the hypersonic deformable aircraft comprises:
selecting an initial attitude motion and aerodynamic model of the hypersonic morphing aircraft based on a geometric model of the hypersonic morphing aircraft, which is expressed as:
let xi= [ alpha beta gamma ] V ] T Is a three-axis attitude angle vector, omega= [ omega ] x ω y ω z ] T For the triaxial angular velocity vector, j=diag (I x ,I y ,I z ) Is an inertia matrix, M= [ M ] x M y M z ] T Is aerodynamic moment, M s ＝[M sx M sy M sz ] T For the deformation additional aerodynamic moment, N is the lateral force, L is the lift, m is the aircraft mass, V is the aircraftSpeed, θ is the velocity tilt angle
Wherein alpha represents attack angle, beta represents sideslip angle, gamma V Representing a roll angle;
the calculation mode for obtaining aerodynamic force and aerodynamic moment of the hypersonic deformation aircraft is expressed as follows:
F i ＝qS 0 C i ,
i＝D,L,N
M i ＝qS 0 L ref C mj ,
j＝x,y,z
wherein κ represents a deformation ratio;
let γ=gf, d 1 ＝H+D 1 ,f(ω)＝J -1 [M s -ω×(Jω)],d 2 ＝D 2 Simplifying the initial gesture motion and the pneumatic model to obtain the gesture motion and the pneumatic model, wherein the gesture motion and the pneumatic model are expressed as follows:
where b is a control matrix, expressed as:
u is a control amount expressed as:
u＝[δ x δ y δ z ] T
wherein delta x Representing roll rudder, delta y Representing yaw rudder, delta z Representing a pitch rudder.
According to one aspect of the invention, in the step S2 of constructing the disturbance observer for estimating the hypersonic deformed aircraft in real time based on the total disturbance in the instruction flight process based on the gesture motion and the aerodynamic model, the disturbance observer is constructed by designing an adaptive law and a convergence criterion and approaching the real disturbance based on a fuzzy logic system.
According to one aspect of the present invention, in step S2, the step of constructing the interference observer by designing an adaptive law and a convergence criterion and approximating the real interference based on the fuzzy logic system includes:
a fuzzy rule form is constructed, which is expressed as:
RULE j:
IF x 1 is
and x n is/>
THEN y is B j
here, the
As fuzzy set, B j And (3) obtaining the fuzzy system output of the anti-fuzzification of the gravity center method for the fuzzy output value of the j-th fuzzy rule, wherein the fuzzy system output is expressed as follows:
here, the
As fuzzy variable x i Membership function of h j Is B j M is the number of fuzzy rules,
to adjust the parameter vector, η (x) = [ η ] 1 (x),η 2 (x),…,η m (x)] T As a fuzzy base function, a fuzzy base function is obtained based on the fuzzy system output, which is expressed as:
building a membership function controlling the fuzzy base function input, which is expressed as:
and constructing an adaptive law and convergence criterion for updating convergence of the weight matrix to finish the construction of the interference observer.
According to one aspect of the present invention, in step S2, in the step of constructing an adaptive law and convergence criterion for updating convergence on the weight matrix, the adaptive law and convergence criterion are expressed as:
according to one aspect of the present invention, in step S3, the step of converting the gesture motion and pneumatic model into a control model facing control, the control model is expressed as:
according to one aspect of the present invention, in step S3, in the step of designing a fixed time-converged sliding mode controller based on a real-time estimation result of the disturbance observer, the sliding mode controller includes: slip form manifold and slip form control law; wherein the slipform manifold is expressed as:
wherein, the liquid crystal display device comprises a liquid crystal display device,
the sliding mode control law adopts a control quantity u of a control model, which is expressed as:
according to the scheme provided by the invention, the problems of rapid stability and tracking of the hypersonic deformed aircraft under the reentry section interference condition are effectively solved, the strong robustness of the control system design method is ensured, the advanced convergence time performance is represented, the requirement of the reentry section on flight tasks is met, and the hypersonic deformed aircraft is applied to the field of aircraft control.
According to one scheme of the hypersonic deformation aircraft, the hypersonic deformation aircraft is designed to fly in a reentry section. Meanwhile, the deformation mode of the deformed aircraft is designed (for example, the deformation mode of the deformed aircraft is selected, namely, the deformed aircraft is stretched), and a pneumatic model is built on the deformed aircraft by using software, wherein the pneumatic model comprises the influence of the deformation on the pneumatic performance. Under the condition of fully considering wing deformation, a reentry section attitude control model suitable for banked turning control is established, and a control method design is carried out based on a fuzzy logic system, an interference observer and a sliding mode manifold, so that a complete and usable hypersonic deformed aircraft reentry section attitude control system design method is formed. The method is suitable for completing the reentry flight task of the hypersonic deformed aircraft, has great significance in engineering application, effectively solves the problems of rapid stability and tracking of the attitude of the hypersonic deformed aircraft under the reentry interference condition, ensures the strong robustness of the design method of the control system, and meets the flight task requirement of the reentry.
According to one aspect of the invention, the invention has fixed time convergence performance for both the estimation of the disturbance and the tracking of the instruction, and the estimated value of the disturbance will be compensated in the sliding mode controller.
According to one scheme of the invention, the interference observer can estimate the system state in real time while estimating the total interference, and the estimation errors of the two can be converged to the vicinity of the origin in fixed time.
According to the scheme provided by the invention, aiming at the complex mode that the required angular velocity is solved from the outer ring angle instruction in the traditional scheme, and then the final rudder deflection is solved by taking the angular velocity as the inner ring instruction, the invention can realize the direct calculation from the outer ring angle instruction to the rudder deflection after the interference estimation error is converged through ingenious model conversion, and the calculated amount is greatly reduced.
According to the scheme of the invention, the fixed-time convergent sliding mode manifold is designed, after the sliding mode manifold reaches and is maintained at the origin, the control model is degraded into an autonomous system, and then the angle tracking error converges to the vicinity of the origin in the fixed time, so that the tracking of the gesture command is completed.
Drawings
FIG. 1 is a block diagram of the steps of a hypersonic morphing aircraft tamper resistant control method according to one embodiment of the invention;
FIG. 2 is a flow chart of a hypersonic morphing aircraft anti-jamming control method according to one embodiment of the present invention;
FIG. 3 is a three-way attitude angle tracking curve for a hypersonic morphing aircraft under nominal conditions according to one embodiment of the present invention;
FIG. 4 is a sliding mode manifold change curve for a hypersonic morphing aircraft under nominal conditions according to one embodiment of the invention;
FIG. 5 is a plot of spanwise deformation rate versus triaxial rudder deflection for a hypersonic deformed aircraft at nominal conditions according to one embodiment of the present invention;
FIG. 6 is a hypersonic morphing aircraft deviation condition with/without disturbance observer command tracking contrast curve according to one embodiment of the present invention;
FIG. 7 is a plot of estimated versus actual values of hypersonic morphing aircraft channel interference according to one embodiment of the present invention;
FIG. 8 is a simulation verification curve of convergence time for a hypersonic morphing aircraft at different initial roll angles according to one embodiment of the invention;
fig. 9 is a plot of attitude angle tracking for a hypersonic morphing aircraft under aerodynamic drag bias conditions according to one embodiment of the present invention.
Detailed Description
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings that are required to be used in the embodiments will be briefly described below.
Referring to fig. 1 and 2, according to an embodiment of the present invention, a hypersonic deformable aircraft anti-interference control method based on fixed time convergence includes:
s1, constructing a gesture motion and pneumatic model based on a geometric model of a hypersonic deformation aircraft;
s2, constructing an interference observer for estimating the hypersonic deformation aircraft in real time based on the total interference in the instruction flight process based on the attitude motion and the pneumatic model;
s3, converting the gesture motion and the pneumatic model into control models facing control, designing a sliding mode controller converging in fixed time based on real-time estimation results of the disturbance observer, and tracking instructions of the hypersonic deformed aircraft by the sliding mode controller based on the control models to complete gesture control tasks under the disturbance condition.
Referring to fig. 1 and 2, in step S1, the step of constructing a pose motion and aerodynamic model based on a geometric model of a hypersonic deformable aircraft according to an embodiment of the present invention includes:
selecting an initial attitude motion and aerodynamic model of the hypersonic deformation aircraft based on a geometric model of the hypersonic deformation aircraft, wherein the initial attitude motion and aerodynamic model is expressed as follows:
wherein alpha represents attack angle, beta represents sideslip angle, gamma V Representing a roll angle;
let xi= [ alpha beta gamma ] V ] T Is a three-axis attitude angle vector, omega= [ omega ] x ω y ω z ] T For the triaxial angular velocity vector, j=diag (I x ,I y ,I z ) Is an inertia matrix, M= [ M ] x M y M z ] T Is aerodynamic moment, M s ＝[M sx M sy M sz ] T For deformation add-on aerodynamic moment, N is lateral force, L is lift, m is aircraft mass, V is aircraft speed, θ is speed pitch, and
the calculation mode for acquiring aerodynamic force and aerodynamic moment of hypersonic deformation aircraft is expressed as:
F i ＝qS 0 C i ,
i＝D,L,N
M i ＝qS 0 L ref C mj ,
j＝x,y,z
wherein, κ represents deformation rate, and the length-variable instruction used for representing the flight of the hypersonic deformation aircraft is independently changed;
let γ=gf, d 1 ＝H+D 1 ,f(ω)＝J -1 [M s -ω×(Jω)],d 2 ＝D 2 The initial gesture motion and the pneumatic model are simplified to obtain the gesture motion and the pneumatic model, which are expressed as:
where b is a control matrix, expressed as:
u is a control amount expressed as:
u＝[δ x δ y δ z ] T
wherein delta x Representing roll rudder, delta y Representing yaw rudder, delta z Representing a pitch rudder.
Referring to fig. 1 and 2, in step S2, in the step of constructing an interference observer for estimating the hypersonic deformed aircraft in real time based on the total interference in the instruction flight process based on the gesture motion and the aerodynamic model, the interference observer is constructed by designing an adaptive law and a convergence criterion and approaching the real interference based on a fuzzy logic system.
As shown in fig. 1, in step S2, the step of constructing the interference observer by designing the adaptive law and the convergence criterion and approximating the real interference based on the fuzzy logic system includes:
a fuzzy rule form is constructed, which is expressed as:
RULE j:
IF x1 is
and xn is/>
THEN y is B j
wherein, the logic expression of the constructed fuzzy rule form is as follows: for rule J, if "x" is entered 1 is
and x n is/>
", then the fuzzy system output is" y is B j ”。
Here, the
As fuzzy set, B j And (3) obtaining the fuzzy system output of the anti-fuzzification of the gravity center method for the fuzzy output value of the j-th fuzzy rule, wherein the fuzzy system output is expressed as follows:
here, the
As fuzzy variable x i Membership function of h j Is B j M is the number of fuzzy rules,
to adjust the parameter vector, η (x) = [ η ] 1 (x),η 2 (x),…,η m (x)] T As a fuzzy base function, a fuzzy base function is obtained based on the fuzzy system output, which is expressed as:
building a membership function controlling the fuzzy base function input, which is expressed as:
in the present embodiment, the state z of the control model obtained in step S3 1 ,z 2 And its differentiation
Is an argument of the membership function and is input in units of degrees.
Constructing an adaptive law and a convergence criterion for updating and converging the weight matrix so as to complete the construction of an interference observer; in this embodiment, the characteristics of the fuzzy logic system that can approach any function are utilized to estimate the interference value, so that the adaptive law and convergence criterion are designed to be used to implement the weight of the fuzzy logic system for estimation and update, so that the system state and total interference of the hypersonic deformed aircraft can be estimated in real time, and the estimation error is at a fixed time T due to the combined action of the higher-order term and the fractional power f1 Internal convergence, while the estimated value of the disturbance will be compensated for in the design of the subsequent sliding mode controller. And has
Wherein->
As shown in fig. 1 and 2, in step S2, in the step of constructing an adaptive law and convergence criterion for updating convergence on the weight matrix, the adaptive law and convergence criterion are expressed as:
as shown in fig. 1 and 2, in the step of converting the gesture motion and pneumatic model into the control model for control in step S3, according to an embodiment of the present invention, based on the gesture motion and pneumatic model obtained in step S1, the gesture motion and pneumatic model are expressed as:
definition of delta 1 ＝g+d 1 ，
B＝Rb，Δ 2 ＝Rd 2 And the state of the transformation model is z 1 ＝ξ-ξ c ，/>
A transformation model, the control model, is obtained, which is expressed as:
through the conversion process, the obtained control model can effectively ensure the solution from the outer ring instruction to the rudder deflection control quantity.
As shown in fig. 1 and 2, in step S3, in the step of designing a sliding mode controller for convergence of a fixed time based on a real-time estimation result of an interference observer, the sliding mode controller includes: slip form manifold and slip form control law; wherein, the slipform manifold is expressed as:
wherein, the liquid crystal display device comprises a liquid crystal display device,
the sliding mode control law adopts the control quantity u of a control model, which is expressed as:
in this embodiment, the sliding mode manifold will also converge at a fixed time T as the disturbance observer converges f2 Inner convergence, and has
The control model then degenerates into an autonomous system, which is represented as:
furthermore, as can be seen from the designed fixed time form and homogeneous system theory, the control model (i.e. autonomous system) will also be at the fixed time T f3 Inner convergence, and has
The whole control model will therefore be at a fixed time T f ＝T f1 +T f2 +T f3 And (5) inner convergence.
The three fixed time convergence characteristics of the disturbance observer, the sliding mode manifold and the autonomous system are contained, and three times T are respectively corresponding f1 、T f2 、T f3 The three are added to form the upper limit of the convergence time of the whole controller.
By the control quantity u, the tracking instruction can be effectively ensured, and meanwhile, the sliding mode manifold is ensured to be converged in a fixed time.
Through the sliding mode manifold, the subsequent autonomous system can be effectively ensured to converge in the fixed time, so that the fixed time convergence design of the whole control system is completed.
According to the invention, both the form of the control quantity u and the expression form of the system of the control model obtained after conversion are calculated by means of the interference estimated value delta obtained by the interference observer, so that the compensation effect of the interference in the control quantity u is effectively realized.
It should be noted that the range of values of the parameters used in the present solution is:
to further illustrate the invention, simulation verification is performed based on an established expert system knowledge base.
In order to verify the effectiveness, anti-interference capability and fixed time convergence characteristics of the invention, numerical simulation is performed. Wherein, the initial state of hypersonic deformation aircraft sets up as: a (0) =0°, β (0) =2°, γ V (0) =0°, κ (0) =1, h=30 km, v=3000 m/s, all initial angular rates being zero. The deformation rate is omega from natural frequency n The second-order link drive with damping ratio of 0.7, which varies between 0 and 1, =10. The simulation step length is 5ms, and the tracking attitude angle instruction is as follows:
the design of the simulation parameters is shown in table 1.
TABLE 1
| Parameters (parameters) | Numerical value | Parameters (parameters) | Numerical value |
| c 1 | [4 1 4] T | γ 1 | 10 4 ×[10 8 8] T |
| c 2 | [2.5 1 5] T | γ 1 ·γ 2 | [0.3 0.3 0.3] T |
| c 3 | 0.001×[1 1 1] T | γ 3 | 10 4 ×[2 1 1] T |
| c 4 | [1 1 1] T | γ 3 ·γ 4 | [0.5 0.5 0.5] T |
| c 5 | [5 5 5] T | r，r 1 ，r 2 | 1.5，0.5，1.5 |
| c 6 | 0.001×[1 1 1] T | μ 1 ，σ 1 | 0.5，1.6 |
| k 1 ，k 2 ，k 3 ，k 4 | 1.1，1.2，9.6，3.5 | l 1 ，l 2 ， |
1，2，0.75 |
2. Analysis of results
The simulation results are shown in fig. 3 to 9 in the drawings of the specification.
Fig. 3-5 are simulation results under nominal conditions, verifying the effectiveness of the present method under nominal conditions. In fig. 3, three-channel attitude angle tracking curves under the nominal condition can be known, and the attack angle, the sideslip angle and the roll angle can all realize rapid and high-precision tracking of instructions. Fig. 4 is a graph of the sliding mode manifold change under nominal conditions, showing that the sliding mode manifold is rapidly converging and stabilized at the origin. Fig. 5 is a graph of the variation of the spread deformation rate and the triaxial rudder deflection under the nominal condition, and it can be seen that the method can still ensure the stable attitude of the aircraft and effectively track the instruction when the aircraft is continuously deformed, and the deflection amplitude and the smoothness of the control rudder are both in a reasonable range.
Fig. 6 to 7 are simulation results under interference conditions, and the estimation effect of the interference observer and the anti-interference performance of the method are verified. Wherein, fig. 6 is a graph showing the tracking contrast of the command of the interference-free observer under the deviation condition, and the control method of the interference-free observer can be used for greatly oscillating the gesture when the interference is applied to the system, so that the control performance is poor, and the method can greatly reduce the negative influence of the interference on the system, realize high-precision gesture control and has extremely strong anti-interference capability. Fig. 7 is a graph comparing the estimated interference value with the actual interference value, and it can be seen that the interference observer can effectively and precisely estimate the interference value in real time in the whole process.
FIG. 8 is a graph of simulated verification of convergence time at different initial roll angles, verifying the fixed time convergence performance of the present method. As can be seen from the figure, taking the roll channel as an example, when the roll angle starts from different directions and different initial magnitudes, and the command angle of 0 degrees is tracked, the convergence time has a fixed upper bound, and the convergence error in this time is strictly zero.
Fig. 9 is a graph of attitude angle tracking under pneumatic bias conditions, verifying the robustness of the method. According to the graph, under different pneumatic combined deflection conditions, the method can realize rapid and high-precision tracking and stabilization of the attitude angle and complete the flight control task.
According to an embodiment of the invention, the hypersonic deformable aircraft comprises a machine body and an onboard circuit board arranged in the machine body. In this embodiment, the onboard circuit board is provided with a processor and a memory, and the memory stores a computer program, and the processor implements the steps of the method when executing the computer program.
The foregoing is merely exemplary of embodiments of the invention and, as regards devices and arrangements not explicitly described in this disclosure, it should be understood that this can be done by general purpose devices and methods known in the art.
The above description is only one embodiment of the present invention, and is not intended to limit the present invention, but various modifications and variations can be made to the present invention by those skilled in the art. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (7)
1. The hypersonic deformed aircraft anti-interference control method based on fixed time convergence is characterized by comprising the following steps of:
s1, constructing a gesture motion and pneumatic model based on a geometric model of a hypersonic deformation aircraft;
s2, constructing an interference observer for estimating the hypersonic deformed aircraft in real time based on total interference in the instruction flight process based on the attitude motion and the pneumatic model;
s3, converting the gesture motion and the pneumatic model into control models facing control, and designing a sliding mode controller converging at fixed time based on real-time estimation results of the disturbance observer, wherein the sliding mode controller tracks instructions of the hypersonic deformation aircraft based on the control models, so that gesture control tasks under disturbance conditions are completed.
2. The hypersonic deformation vehicle anti-interference control method according to claim 1, wherein in step S1, the step of constructing the attitude motion and aerodynamic model based on the geometric model of the hypersonic deformation vehicle includes:
selecting an initial attitude motion and aerodynamic model of the hypersonic morphing aircraft based on a geometric model of the hypersonic morphing aircraft, which is expressed as:
let xi= [ aβγ ] V ] T Is a three-axis attitude angle vector, omega= [ omega ] x ω y ω z ] T For the triaxial angular velocity vector, j=diag (I x ,I y ,I z ) Is an inertia matrix, M= [ M ] x M y M z ] T Is aerodynamic moment, M s ＝[M sx M sy M sz ] T For deformation add-on aerodynamic moment, N is lateral force, L is lift, m is aircraft mass, V is aircraft speed, θ is speed pitch, and
wherein alpha represents attack angle, beta represents sideslip angle, gamma V Representing a roll angle;
the calculation mode for obtaining aerodynamic force and aerodynamic moment of the hypersonic deformation aircraft is expressed as follows:
F i ＝qS 0 C i ,
M i ＝qS 0 L ref C mj ,
wherein κ represents a deformation ratio;
let γ=gf, d 1 ＝H+D 1 ,f(ω)＝J -1 [M s -ω×(Jω)],d 2 ＝D 2 Simplifying the initial gesture motion and the pneumatic model to obtain the gesture motion and the pneumatic model, wherein the gesture motion and the pneumatic model are expressed as follows:
where b is a control matrix, expressed as:
u is a control amount expressed as:
u＝[δ x δ y δ z ] T
wherein delta x Representing roll rudder, delta y Representing yaw rudder, delta z Representing a pitch rudder.
3. The hypersonic deformation aircraft anti-interference control method according to claim 2, wherein in the step of constructing an interference observer for estimating the hypersonic deformation aircraft in real time based on total interference in the instruction flight process based on the gesture motion and the aerodynamic model in the step S2, the interference observer is constructed by designing an adaptive law and a convergence criterion and approaching a real interference based on a fuzzy logic system.
4. The hypersonic deformation vehicle anti-interference control method according to claim 3, wherein in step S2, the step of constructing the interference observer by designing an adaptive law and a convergence criterion and approaching a real interference based on a fuzzy logic system includes:
a fuzzy rule form is constructed, which is expressed as:
RULE j:
THEN y is B j
here, the
As fuzzy set, B j Obtaining the anti-fuzzification of the gravity center method for the fuzzy output value of the j-th fuzzy ruleIs represented as:
here, the
As fuzzy variable x i Membership function of h j Is B j M is the number of fuzzy rules,
to adjust the parameter vector, η (x) = [ η ] 1 (x),η 2 (x),…,η m (x)] T As a fuzzy base function, a fuzzy base function is obtained based on the fuzzy system output, which is expressed as:
building a membership function controlling the fuzzy base function input, which is expressed as:
and constructing an adaptive law and convergence criterion for updating convergence of the weight matrix to finish the construction of the interference observer.
5. The hypersonic deforming aircraft anti-interference control method according to claim 4, wherein in the step S2, the step of constructing an adaptive law and convergence criterion for updating convergence of the weight matrix is expressed as:
6. the hypersonic deforming aircraft anti-jamming control method as recited in claim 5, wherein in the step of converting the attitude motion and aerodynamic model into a control-oriented control model in step S3, the control model is expressed as:
7. the hypersonic deforming aircraft anti-jamming control method according to claim 6, wherein in the step of designing a fixed time converging sliding mode controller based on the real-time estimation result of the jamming observer in step S3, the sliding mode controller includes: slip form manifold and slip form control law; wherein the slipform manifold is expressed as:
wherein, the liquid crystal display device comprises a liquid crystal display device,
the sliding mode control law adopts a control quantity u of a control model, which is expressed as:
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