CN113885552B

CN113885552B - Preset performance control method and system for hypersonic aircraft

Info

Publication number: CN113885552B
Application number: CN202111021060.8A
Authority: CN
Inventors: 李海燕; 韦俊宝; 李静; 方登建; 袁胜智; 胡云安; 王斌; 吴佳栋; 高飞
Original assignee: Naval University of Engineering PLA
Current assignee: Naval University of Engineering PLA
Priority date: 2021-09-01
Filing date: 2021-09-01
Publication date: 2023-09-29
Anticipated expiration: 2041-09-01
Also published as: CN113885552A

Abstract

The invention provides a preset performance control method and a system for a hypersonic aircraft, wherein the method comprises the following steps: based on tracking errors, constructing a preset performance function of the hypersonic aircraft; constructing a speed subsystem controller according to a preset performance function and a saturation function; according to a preset performance function and a limited instruction filter, constructing a height subsystem controller by an inversion control method; acquiring a state initial value of the hypersonic aircraft at the current moment, and performing tracking control according to a speed subsystem controller and a height subsystem controller; the saturation function is constructed according to the fuel equivalence ratio and the elevator deflection angle; the constrained-instruction filter is constructed based on the input saturation problem from the elevator deflection angle ideal control input value. On the basis of improving the steady state and transient performance of the system, the invention ensures that the output tracking error has smaller overshoot; the amplitude and the speed of the system input are ensured to meet the limited requirement, and good tracking performance is provided.

Description

Preset performance control method and system for hypersonic aircraft

Technical Field

The invention relates to the technical field of automatic control, in particular to a preset performance control method and system for a hypersonic aircraft.

Background

Hypersonic aircraft is a novel aircraft flying at speeds above Mach 5 in near space, and has great application potential in civil and military fields. Currently, research on hypersonic aircraft control technology has achieved a certain result. The hypersonic aircraft has high requirements on the dynamic performance of the control system when flying at high speed. The preset performance method has the unique advantage of simultaneously considering transient and steady performance of the system, and is widely applied to hypersonic aircraft control research.

In practical control systems, the control force provided by the actuator is limited. The problem of saturation of the output of the actuator is easily caused by the influence of the high-altitude flight of the aircraft and the external environment. Once the system is saturated, the ideal control law cannot be effectively executed, so that larger deviation of instruction tracking can occur, and even the stability of the system is seriously affected.

Accordingly, there is a need for a method and system for controlling the performance of hypersonic aircraft.

Disclosure of Invention

Aiming at the problems existing in the prior art, the invention provides a preset performance control method and system for a hypersonic aircraft.

The invention provides a preset performance control method for a hypersonic aircraft, which comprises the following steps:

based on tracking errors, constructing a preset performance function of the hypersonic aircraft;

constructing a speed subsystem controller according to the preset performance function and the saturation function;

constructing a height subsystem controller through an inversion control method according to the preset performance function and the limited instruction filter;

acquiring a state initial value of the hypersonic aircraft at the current moment, and tracking and controlling the hypersonic aircraft according to the speed subsystem controller and the altitude subsystem controller;

the saturation function is constructed according to the fuel oil equivalent ratio and the deflection angle of the elevator; the limited instruction filter is constructed based on the input saturation problem and according to the ideal control input value of the elevator deflection angle.

According to the preset performance control method for the hypersonic aircraft, which is provided by the invention, the preset performance function of the hypersonic aircraft is constructed based on tracking errors, and the preset performance control method comprises the following steps:

aiming at minimizing the overshoot of the tracking error of the hypersonic aircraft, constructing a preset performance function of the hypersonic aircraft, wherein the preset performance function is as follows:

p ₂ (t)＜e(t)＜p ₁ (t)；

wherein ,p₁(t) and p₂ (t) represents a preset performance function, e (0) represents a tracking error at an initial time,representing an existing performance function; />And mu is constant, < >>Is a steady state value +.>And->μ＞0。

According to the preset performance control method for the hypersonic aircraft provided by the invention, a speed subsystem controller is constructed according to the preset performance function and the saturation function, and the method comprises the following steps:

according to the preset performance function, constructing a speed performance function of the hypersonic aircraft, wherein the speed performance function is as follows:

υ _V ＝e _V -ξ _V ；

e _V ＝V-V _d ；

p _V2 ＜υ _V ＜p _V1 ；

wherein ,p_V1(t) and p_V2 (t) represents a preset performance function established for the aircraft speed V, V _V Representing speed compensation errors, e _V Representing velocity tracking error, ζ _V Representing the auxiliary variable to be designed, V _d Representing a speed command, V representing the aircraft speed; mu (mu) _V 、 and />Is a speed performance function parameter; sigma (sigma) _V > 0, constant;

performing error transformation on the speed performance function to obtain a speed error transformation function of the hypersonic aircraft, wherein the speed error transformation function is as follows:

wherein ,ε_V Representing a speed transformation error;

based on the hypersonic aircraft longitudinal motion rigid body model, the velocity subsystem model is obtained as follows:

wherein Φ represents the fuel equivalence ratio, d ₁ The first disturbance term is represented by a, the attack angle is represented by D, the resistance is represented by g, the gravitational acceleration is represented by gamma, the track inclination angle is represented by m, the mass is represented by T ₀(α) and T_Φ (α) represents a thrust-related aerodynamic parameter;

according to the speed subsystem model, deriving the speed error transformation function, and constructing a speed subsystem control law according to the derived speed error transformation function:

wherein ,Φ_d Represents the ideal control input value, k, of the fuel equivalence ratio _V and λ_V Is a positive parameter, the parameter is a positive parameter,representing the first derivative of the speed command, +.>Representing the first derivative, k, of an existing performance function parameter _Vξ Representing auxiliary system parameters->Representing an estimate of the first interference term;

based on the fuel equivalence ratio saturation function, constructing a speed subsystem controller according to the speed subsystem control law:

the fuel equivalence ratio saturation function is:

wherein the constant phi _max and Φ_min The upper and lower limits of the amplitude of the fuel equivalence ratio phi are respectively set.

According to the preset performance control method for the hypersonic aircraft provided by the invention, before the speed subsystem controller is constructed according to the speed subsystem control law based on the fuel equivalent ratio saturation function, the method further comprises:

according to an ideal control input value of the fuel equivalent ratio and an actual input value of the fuel equivalent ratio, a first auxiliary system is constructed and used for ensuring stable tracking when the fuel equivalent ratio is saturated, and the first auxiliary system is as follows:

Where Φ represents the actual input value of the fuel equivalence ratio.

According to the preset performance control method for the hypersonic aircraft, according to the preset performance function and the limited instruction filter, an altitude subsystem controller is constructed through an inversion control method, and the method comprises the following steps:

according to the preset performance function, constructing a high performance function of the hypersonic aircraft, wherein the high performance function is as follows:

e _h ＝h-h _d ；

p _h2 ＜e _h ＜p _h1 ；

wherein ,p_h1(t) and p_h2 (t) represents a preset performance function built for the altitude h of the aircraft, e _h Representing altitude error, h represents aircraft altitude, h _d Indicates the altitude command, mu _h 、 and />Is a high performance function parameter; sigma (sigma) _h > 0, constant;

performing error transformation on the altitude performance function to obtain an altitude error transformation function of the hypersonic aircraft, wherein the altitude error transformation function is as follows:

wherein ,ε_h Representing a height conversion error;

based on the hypersonic aircraft longitudinal motion rigid body model, the altitude subsystem model is obtained as follows:

wherein V represents the speed of the aircraft, gamma represents the track inclination angle, theta represents the pitch angle, q represents the pitch angle speed, and d ₂ Representing a second interference term, d ₃ Representing a third interference term, delta _e Represents the deflection angle of the elevator, L ₀ and L_α Representing lift-related aerodynamic parameters, g representing gravitational acceleration, M representing mass, M _T 、M ₀(α) and Representing the relevant parameters of pitching moment, I _yy Representing moment of inertia;

based on the altitude instruction, constructing a first virtual control law gamma _d According to the first virtual control law gamma _d Define the track inclination error as e _γ ＝γ-γ _d And according to the altitude subsystem model, deriving the track inclination angle error to obtain the derived track inclination angle error:

based on the track inclination angle error, a second virtual control law is constructed, and the track inclination angle control law is constructed according to the second virtual control law and the track inclination angle error after derivation:

wherein ,k_γ > 0, which is the relevant parameter of the track dip angle;is d ₂ Estimated value of e _γ Representing track pitch error, θ _d Is χ as the second virtual control law _γ2 Representing the first derivative +.>Is a function of the estimated value of (2);

according to the second virtual control law theta _d Define pitch angle error as e _θ ＝θ-θ _d And according to the altitude subsystem model, deriving the pitch angle error to obtain the derived pitch angle error:

based on the pitch angle error and the track dip angle error, a third virtual control law is constructed, and a pitch angle control law is constructed according to the third virtual control law and the pitch angle error after derivation:

wherein ,k_θ > 0, the relevant design parameters for pitch angle; zeta type toy _q Is an auxiliary variable to be designed; x is x _θ2 For the second virtual control law derivativeIs a function of the estimated value of (2);

defining pitch angle speed error as e _q ＝q-q _d The pitch angle compensation error is v _q ＝e _q -ξ _q According to the altitude subsystem model, deriving the pitch angle compensation error to obtain a derived pitch angle compensation error:

wherein ,q_d A pitch angle rate command;

based on the elevator deflection angle saturation function, constructing a limited instruction filter according to an elevator deflection angle ideal control input value, and obtaining an elevator deflection angle actual input value:

the elevator deflection angle saturation function is as follows:

wherein ,τ_δ and ω_δ Is positive parameter, delta _ed For ideal control input value of elevator deflection angle, constant delta _max and δ_min Respectively the deflection angles delta of elevators _e Upper and lower limits of amplitude;is the derivative of the elevator deflection angle; />Constant psi _max and ψ_min Respectively the deflection angles delta of elevators _e Upper and lower limits of the rate;

constructing a second auxiliary system for counteracting the influence of input saturation according to the ideal control input value of the elevator deflection angle and the actual input value of the elevator deflection angle:

constructing an elevator deflection angle control law:

wherein ,k_q > 0, a relevant design parameter for pitch angle rate; For the third interference term d ₃ Is χ _q2 Derivative of the third virtual control law +.>Estimate of k _qξ1 and k_qξ2 Is an auxiliary system parameter;

constructing a pitch angle speed compensation error control law according to the second auxiliary system, the elevator deflection angle control law and the derived pitch angle compensation error:

and constructing an altitude subsystem controller according to the track inclination angle control law, the pitch angle control law and the pitch angle speed compensation error control law.

According to the preset performance control method for the hypersonic aircraft provided by the invention, before the state initial value of the hypersonic aircraft at the current moment is obtained and the hypersonic aircraft is tracked and controlled according to the speed subsystem controller and the altitude subsystem controller, the method further comprises:

based on interference existing in the hypersonic aircraft during operation, a second-order linear extended state observer is constructed and used for observing and compensating the interference, and the formula of the second-order linear extended state observer is as follows:

wherein ,for the estimated value of the aircraft speed V +.>Is the estimated value of track dip angle gamma, +.>Is an estimated value of pitch angle rate q; / >As interference term d _i (i=1, 2, 3); l (L) _V1 ,l _V2 ,l _γ1 ,l _γ2 ,l _q1 ,l _q2 Are all positive parameters omega ₀ Representing the bandwidth of the observer, parameter a _i =3-! I-! (3-i) ≡! (i=1, 2); phi represents the fuel equivalent ratio, theta represents the pitch angle, q represents the pitch angle rate, gamma represents the track pitch angle, delta _e Indicating the elevator deflection angle.

The invention also provides a preset performance control system for a hypersonic aircraft, comprising:

the performance function construction module is used for constructing a preset performance function of the hypersonic aircraft based on the tracking error;

the speed subsystem controller construction module is used for constructing a speed subsystem controller according to the preset performance function and the saturation function;

the height subsystem controller construction module is used for constructing a height subsystem controller through an inversion control method according to the preset performance function and the limited instruction filter;

the preset performance control module is used for acquiring a state initial value of the hypersonic aircraft at the current moment and carrying out tracking control on the hypersonic aircraft according to the speed subsystem controller and the altitude subsystem controller;

The invention also provides an electronic device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, the processor implementing the steps of the preset performance control method for a hypersonic aircraft as described in any one of the above when executing the program.

The invention also provides a non-transitory computer readable storage medium having stored thereon a computer program which, when executed by a processor, implements the steps of a preset performance control method for a hypersonic aircraft as described in any one of the above.

The invention also provides a computer program product comprising a computer program which, when executed by a processor, implements the steps of a preset performance control method for a hypersonic aircraft as described in any one of the above.

According to the preset performance control method and system for the hypersonic aircraft, the new preset performance function is designed, so that the output tracking error has smaller overshoot on the basis of improving the steady state and transient performance of the hypersonic aircraft system; meanwhile, by constructing the limited instruction filter, the amplitude and the speed of the system input are ensured to meet the limited requirement, so that good tracking performance can be provided on the basis of solving the problem of limited amplitude and speed of the system input.

Drawings

In order to more clearly illustrate the invention or the technical solutions of the prior art, the following description will briefly explain the drawings used in the embodiments or the description of the prior art, and it is obvious that the drawings in the following description are some embodiments of the invention, and other drawings can be obtained according to the drawings without inventive effort for a person skilled in the art.

FIG. 1 is a schematic flow chart of a preset performance control method for a hypersonic aircraft provided by the invention;

FIG. 2 is a schematic diagram of an inequality constraint curve of a preset performance provided by the present invention;

FIG. 3 is a schematic diagram of an inequality constraint curve of a preset performance function provided by the present invention;

FIG. 4 is a comparative schematic diagram of a velocity and tracking error curve provided by the present invention;

FIG. 5 is a schematic diagram showing the comparison of the height and tracking error curves provided by the present invention;

FIG. 6 is a comparative schematic of a system state variable curve provided by the present invention;

FIG. 7 is a schematic diagram showing a comparison of tracking error curves of system state variables according to the present invention;

FIG. 8 is a graphical representation of a comparison of fuel equivalence ratio curves provided by the present invention;

FIG. 9 is a comparative schematic of an elevator deflection angle curve provided by the present invention;

FIG. 10 is a graphical representation of a comparison of elevator yaw rate curves provided by the present invention;

FIG. 11 is a comparative schematic of an auxiliary variable curve provided by the present invention;

FIG. 12 is a comparative schematic of LESO observation curves provided by the present invention;

FIG. 13 is a schematic diagram of a configuration of a preset performance control system for a hypersonic vehicle in accordance with the present invention;

fig. 14 is a schematic structural diagram of an electronic device provided by the present invention.

Detailed Description

For the purpose of making the objects, technical solutions and advantages of the present invention more apparent, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings, and it is apparent that the described embodiments are some embodiments of the present invention, not all embodiments. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.

According to the hypersonic aircraft longitudinal motion rigid model, the preset performance control of the hypersonic aircraft is decomposed into the speed subsystem and the altitude subsystem, so that the preset performance control of the hypersonic aircraft is realized according to the speed subsystem and the altitude subsystem, and the hypersonic aircraft longitudinal motion rigid model is as follows:

Wherein, the aircraft speed V, the aircraft height h, the track dip angle gamma, the pitch angle theta and the pitch angle speed q are rigid state variables; α is the angle of attack and has α=θ - γ; m is mass, g is gravitational acceleration, I _yy Is the moment of inertia; t, D, L, M are thrust, drag, lift and pitching moment respectively, and the corresponding formulas can be described as:

wherein q=0.5 ρv ² Dynamic pressure of the aircraft, wherein ρ is air density; s is the reference area of the aircraft, phi is the fuel equivalence ratio, delta _e Is the elevator deflection angle; and />Is the relevant pneumatic parameter of resistance, L ₀ and L_α Is the relevant aerodynamic parameter of lift force, T _Φ(α) and T₀ (α) is a thrust-related aerodynamic parameter; m is M _T ，M ₀(α) and />is a relevant parameter of the pitching moment.

In the hypersonic aircraft longitudinal motion rigid body model, the Tsin alpha term value in the formula 3 is assumed to be far smaller than the lift force L value, so that the term can be ignored, and the invention takes the term as the assumption 1.

Further, the outputs of the system model are the aircraft speed V and the aircraft altitude h; the control inputs are the fuel equivalence ratio phi and the elevator deflection angle delta _e . Equations 1 through 5, in combination with the hypersonic aircraft longitudinal motion rigid body model, and assuming 1, it is known that the variation of the aircraft velocity V is mainly controlled by the fuel equivalence ratio Φ; elevator deflection delta _e The change of the pitch angle theta and the track dip angle gamma is controlled by directly controlling the change of the pitch angle rate q, so that the change of the aircraft height h is mainly influenced by the deflection angle delta of the elevator _e Is controlled by the control system. In order to facilitate control law design, the invention is based on formulas 1 to 5 of the hypersonic aircraft longitudinal motion rigid body model, and can be decomposed into a speed subsystem model and a height subsystem model:

wherein ,d_i (i=1, 2, 3) is an interference term, comprising external interference and parametric perturbation, for which the invention defines hypothesis 2: interference term d _i (i=1, 2, 3) is continuous and the first derivative is bounded.

Furthermore, in order to avoid the thermal resistance phenomenon of the hypersonic aircraft in actual flight, the fuel equivalent ratio phi value is required to be in a certain range, so that the scramjet engine always maintains a reasonable working state, and otherwise, the engine stops working. In addition, since the actual physical mechanism has a deflection limit, the elevator deflects an angle delta _e The output is also limited. The invention is therefore based on the fuel equivalence ratio phi and the elevator deflection delta _e Under the limited condition, the corresponding saturation function is constructed and can be respectively described as:

wherein ,Φ_d Ideal control input value for fuel equivalence ratio, delta _ed For ideal control input value of elevator deflection angle, constant phi _max and Φ_min Respectively the magnitude of the fuel equivalent ratio phi (i.e. the actual input of the fuel equivalent ratio phiValue), constant delta _max and δ_min Respectively the deflection angles delta of elevators _e Amplitude (i.e. elevator deflection delta _e Actual input value);is the derivative of the elevator deflection angle; />Constant psi _max and ψ_min Respectively the deflection angles delta of elevators _e Upper and lower limits of the rate.

Fig. 1 is a schematic flow chart of a preset performance control method for a hypersonic aircraft, provided by the invention, as shown in fig. 1, and the preset performance control method for the hypersonic aircraft is characterized by comprising the following steps:

step 101, constructing a preset performance function of the hypersonic aircraft based on the tracking error.

The constraint of the existing performance function can ensure that the tracking error has a smaller steady state value, has a certain effect on the improvement of the transient performance of the error, and still can possibly cause the problem of overlarge overshoot of the tracking error. Based on the defects of the existing preset performance method, the method aims at minimizing the overshoot of the tracking error, constructs a new preset performance function, enables the overshoot of the tracking error to be smaller on the basis of considering the transient and steady performance of the system, and improves the dynamic performance of the tracking error of the system.

Specifically, in the existing preset performance control method, tracking errors are limited to a preset convergence region, so that the system meets preset transient and steady-state performance requirements. The constraint inequality of the existing preset performance to the error is as follows:

wherein sigma is a constant, and 0 < sigma is less than or equal to 1;the expression is as follows, which is the existing performance function:

wherein ,μ and />Are all constant and->μ＞0，/>At the time of the steady-state value,from the existing formulas for performance functions described above, the performance functions have a continuously bounded, monotonically decreasing nature.

Fig. 2 is a schematic diagram of an inequality constraint curve of the existing preset performance provided by the present invention, and referring to fig. 2, although the constraint of the existing performance function can ensure a steady state value with a smaller tracking error and has a certain effect on improving the transient performance of the error, the problem of excessive overshoot of the tracking error may still occur. Therefore, for improving the transient performance of tracking error, the existing preset performance method has certain defects. Based on the above problems, the present invention designs a new performance function, specifically, the step 101 includes:

It is obvious that the process is not limited to,further, the inequality constraint is set as:

p ₂ (t)＜e(t)＜p ₁ (t); formula (15)

FIG. 3 is a schematic diagram of an inequality constraint curve of a preset performance function according to the present invention, as shown in FIG. 3, taking e (0) > 0 as an example, by designing parameters such that p ₂ (0) > 0, the maximum deviation of tracking error is smaller thanError steady state value is +.>Within the range. Thus, the present invention provides a preset performance functionCompared with the existing performance function, the error tracking precision is improved, and meanwhile, the error can be ensured to have smaller overshoot. Note that, when the initial error is zero, the inequality (15) is still true, i.e., when e (0) =0, then +.>Inequality (15) becomes +.>

Because the controller is difficult to design by directly utilizing the inequality (15), the inequality constraint is converted into the equality constraint, and the error conversion is carried out, so that the method comprises the following steps of:

based on the above formula, the present invention sets theorem 1: if ε (t) is bounded, then inequality (15) holds, i.e., the system tracking error is not only bounded but also limited to a preset range.

Further, the above theorem 1 is proved to be specifically:

equation (17) can be transformed into:

further, the method comprises the steps of,

since ε (t) is bounded, thenCombining equation (16) and equation (19), one can obtain:

thus, p ₂ (t)＜e(t)＜p ₁ (t)。

And 102, constructing a speed subsystem controller according to the preset performance function and the saturation function.

In the present invention, the control targets are: under the condition that the hypersonic aircraft control system is considered to have limited input, the system output can stably track the command signal, and the amplitude and the speed of the actuating mechanism meet the limit requirement; and after the saturation of the system is finished, the tracking error output by the system reaches the preset requirements of transient and steady-state performances. The step 102 specifically includes:

/>

e _V ＝V-V _d the method comprises the steps of carrying out a first treatment on the surface of the Formula (23)

υ _V ＝e _V -ξ _V The method comprises the steps of carrying out a first treatment on the surface of the Formula (24)

p _V2 ＜υ _V ＜p _V1 The method comprises the steps of carrying out a first treatment on the surface of the Formula (25)

wherein ,p_V1(t) and p_V2 (t) represents a preset performance function established for the aircraft speed V, V _V Representing speed compensation errors, e _V Representing velocity tracking error, ζ _V Representing the auxiliary variable to be designed, V _d Representing a speed command, V representing the aircraft speed; mu (mu) _V 、 and />Is a speed performance function parameter; sigma (sigma) _V > 0, constant. In the invention, the tracking error about the speed of the aircraft is defined by the formula (7), namely the formula (23), and further the compensation error is defined, namely the formula (24). In compensating error v for speed _V A constraint inequality (i.e., equation 25) is established.

wherein ,ε_V Representing a speed transformation error;

wherein f is set to _V and g_V As a substitute symbol, reference is conveniently made to the two formulas above; phi represents the fuel equivalence ratio, d ₁ Representing a first interference term, a representing an angle of attackD represents resistance, g represents gravitational acceleration, gamma represents track inclination, m represents mass, T ₀(α) and T_Φ (α) represents a thrust-related aerodynamic parameter;

deriving the speed error transformation function according to the speed subsystem model, specifically, combining formula (7) and deriving formula (26), thereby obtaining:

and constructing a speed subsystem control law according to the derived speed error transformation function:

/>

wherein ,Φ_d Represents the ideal control input value, k, of the fuel equivalence ratio _V and λ_V Is a positive parameter, the parameter is a positive parameter,representing the first derivative of the speed command; similarly, v is set _V and r_V As a substitute symbol, reference is conveniently made to the two formulas above; />Representing the first derivative, k, of an existing performance function parameter _Vξ Representing auxiliary system parameters->Representing an estimate of the first interference term;

based on the fuel equivalence ratio saturation function, taking into account the input saturation problem, a formula (9) in the saturation function is introduced for constraining the ideal control input Φ _d And get the actual input:

Φ＝H _Φ (Φ _d ) The method comprises the steps of carrying out a first treatment on the surface of the Formula (29)

And constructing a speed subsystem controller according to the speed subsystem control law, namely substituting the formula (28) and the formula (29) into the formula (27) to obtain:

the fuel equivalence ratio saturation function is:

On the basis of the above embodiment, before the speed subsystem controller is constructed according to the speed subsystem control law based on the fuel equivalence ratio saturation function, the method further includes:

Where Φ represents the actual input value of the fuel equivalence ratio. By designing the first auxiliary system, stable tracking of the system when the fuel oil equivalent ratio phi is saturated is ensured.

And step 103, constructing a height subsystem controller through an inversion control method according to the preset performance function and the limited instruction filter. Step 103 specifically includes:

e _h ＝h-h _d ；

p _h2 ＜e _h ＜p _h1 the method comprises the steps of carrying out a first treatment on the surface of the Formula (33)

wherein ,ε_h Representing a high conversion error.

wherein V represents the speed of the aircraft, and gamma represents the track inclination angle; for convenience of reference, f _γ 、g _γ 、f _q and g_q As a surrogate for the above formulas; θ represents pitch angle, q represents pitch angle rate, d ₂ Representing a second interference term, d ₃ Representing a third interference term, delta _e Represents the deflection angle of the elevator, L ₀ and L_α Representing lift-related aerodynamic parameters, g representing gravitational acceleration, M representing mass, M _T 、M ₀(α) and Representing the relevant parameters of pitching moment, I _yy Representing moment of inertia;

based on the altitude instruction, constructing a first virtual control law gamma _d . In the present invention, in order to achieve the altitude command h for the aircraft altitude h _d Is designed as follows, i.e. a first virtual control law:

wherein ,k_h > 0, is a parameter. In the present inventionIn the method, when the track dip angle gamma is implemented to gamma _d Conversion error epsilon during tracking of (a) _h (t) satisfyI.e. epsilon _h (t) is bounded. Therefore, according to the above theorem 1, it is known that when γ→γ _d Height tracking error e _h The preset transient and steady state performance requirements are met.

Further, a height subsystem control law is designed by an inversion control method.

According to the first virtual control law gamma _d Define the track inclination error as e _γ ＝γ-γ _d And deriving the track inclination angle error according to the altitude subsystem model, namely by combining the formula (8), so as to obtain the derived track inclination angle error:

in the conventional instruction filter, the following is specifically mentioned:

wherein ,x_d Is a filter input; output χ ₁ ,χ ₂ Respectively x _d X is a group _d First derivativeIs a function of the estimated value of (2); τ, ω are filter parameters, and τ e (0, 1)]，ω＞0，/> The present invention therefore defines hypothesis 3: the presence of an unknown constant eta ₁ ＞0,η ₂ > 0, such that% ₁ -x _d |≤η ₁ ，/>In the invention, the virtual command gamma can be estimated by using a command filter in consideration of the difficulty in deriving the virtual command in the design of the inversion controller _d And its derivative->

wherein ,τ_γ ,ω _γ Are positive parameters.

Then, based on the track pitch error, a second virtual control law is constructed:

wherein ,k_γ The values of > 0 are parameters,is d ₂ And (5) estimating a value.

And substituting the formula (39) into the formula (36) according to the second virtual control law and the derived track dip angle error to construct the track dip angle control law:

according to the second virtual control law theta _d Define pitch angle error as e _θ ＝θ-θ _d And according to the altitude subsystem model, namely the formula (8), deriving the pitch angle error to obtain the derived pitch angle error:

based on the pitch angle error and the track inclination angle error, a third virtual control law is constructed, specifically:

q _d ＝-k _θ e _θ +g _γ e _γ -ξ _q +χ _θ2 The method comprises the steps of carrying out a first treatment on the surface of the Formula (42)

wherein ,k_θ > 0 is a design parameter, ζ _q Is an auxiliary variable to be designed; x-shaped articles _θ2 Is the second one

Derivative of virtual control lawCan be obtained by the following instruction filter:

wherein ,τ_θ ,ω _θ Are positive parameters.

And substituting the formula (42) into the formula (41) according to the third virtual control law and the pitch angle error after derivation, and constructing to obtain a pitch angle control law:

wherein ,k_θ > 0, the relevant design parameters for pitch angle; zeta type toy _q Is an auxiliary variable to be designed; x-shaped articles _θ2 For the second virtual control law derivativeIs a function of the estimated value of (2);

further, defining a pitch angle rate error as e _q ＝q-q _d The pitch angle compensation error is v _q ＝e _q -ξ _q And according to the altitude subsystem model, namely combining the formula (8), deriving the pitch angle compensation error to obtain the derived pitch angle compensation error:

wherein ,q_d A pitch angle rate command;

based on the elevator deflection angle saturation function, a limited instruction filter is constructed according to an elevator deflection angle ideal control input value, and an elevator deflection angle actual input value is obtained. In the present invention, a constrained instruction filter is constructed to constrain the ideal control input delta in consideration of the input saturation problem _ed And get the actual input:

The elevator deflection angle saturation function is as follows:

/>

wherein ,τ_δ and ω_δ Is positive parameter, delta _ed For ideal control input value of elevator deflection angle, constant delta _max and δ_min Respectively the deflection angles delta of elevators _e Upper and lower limits of amplitude;is the derivative of the elevator deflection angle; />Constant psi _max and ψ_min Respectively the deflection angles delta of elevators _e Upper and lower limits of the rate. By constructing the limited instruction filter, the ideal control law delta for the deflection angle of the elevator is input into the filter _ed The output is the actual control law delta of the deflection angle of the elevator _e The filter functions to make the actual control law delta _e Meeting the limited requirements of amplitude and speed.

In order to offset the influence caused by input saturation, the invention constructs a second auxiliary system for offset the influence of input saturation according to the ideal control input value of the elevator deflection angle and the actual input value of the elevator deflection angle:

constructing an elevator deflection angle control law:

wherein ,k_q > 0, a relevant design parameter for pitch angle rate;for the third interference term d ₃ Is χ _q2 Derivative of the third virtual control law +.>Estimate of k _qξ1 and k_qξ2 Is an auxiliary system parameter. In the present invention, χ _q2 The method can be obtained by the following instruction filter:

wherein ,τ_q ,ω _q Are positive parameters.

According to the second auxiliary system, the elevator deflection angle control law and the derived pitch angle compensation error, substituting a formula (47) and a formula (48) into a formula (45), and constructing a pitch angle speed compensation error control law:

and finally, constructing a height subsystem controller according to the track inclination angle control law, the pitch angle control law and the pitch angle speed compensation error control law. And therefore, the altitude subsystem controller is constructed to track and control the aircraft.

104, acquiring a state initial value of the hypersonic aircraft at the current moment, and tracking and controlling the hypersonic aircraft according to the speed subsystem controller and the altitude subsystem controller;

the saturation function is constructed according to the fuel equivalent ratio and the elevator deflection angle.

In the present invention, theorem 2 is set: for formulas (1) to (5) of the hypersonic aircraft longitudinal motion rigid body model, the fuel equivalent ratio phi and the elevator deflection angle delta can be ensured by adopting the formulas (29) and (46) to restrain the system input _e The actual output of (2) always meets the limited condition, i.e. Φ e [ Φ ] _min ,Φ _max ]，δ _e ∈[δ _min ,δ _max ]，

Further, the attestation theorem 2 is performed by: since Φ=h _Φ (Φ _d )，δ _e ＝H _δ (χ _δ1 ) According to the saturation function H _Φ(·) and H_δ Definition of (-), getTo phi epsilon phi _min ,Φ _max ]，δ _e ∈[δ _min ,δ _max ]。

For the constrained instruction filter constructed by the present invention, equation (46), the equation is then appliedAnd (3) performing transformation to obtain:

wherein ,c_δ ＝2τ _δ ω _δ . Due to saturation functionThen further result according to equation (51):

multiplying the inequality, equation (52), by exp (c) _δ t), obtaining:

c _δ ψ _min exp(c _δ t)≤(χ _δ2 exp(c _δ t))′≤c _δ ψ _max exp(c _δ t); formula (53)

Further, the inequality (53) is integrated to obtain:

according to actual conditions, the elevator deflection angle delta _e Upper and lower limits ψ of output rate _min ＜0,ψ _max > 0. Take the initial value χ _δ2 (0) =0, then equation (54) can be reduced to:

ψ _min ≤χ _δ2 ≤ψ _max the method comprises the steps of carrying out a first treatment on the surface of the Formula (55)

When χ is _δ1 ∈(δ _min ,δ _max ) When then delta _e ＝H _δ (χ _δ1 )＝χ _δ1 ObviouslyWhen (when)δ _e (t)＝δ _min Or delta _e (t)＝δ _max Then->The method is available in a comprehensive way,it should be noted that, in the prior art, with respect to the control input amplitude and rate limiting problem, the limiting instruction filter is constructed as follows:

if the function H is as shown in equation (56) _ψ (. Cndot.) reaches saturation, at which point equation (56) becomes:

wherein the constant psi _m ＝ψ _min Or psi _m ＝ψ _max . Obviously, equation (57) does not guarantee the output δ _e The amplitude of (a) satisfies the constraint. Therefore, the limited instruction filter constructed by the prior art cannot guarantee that effective constraint on control input can be achieved.

Further, the present invention sets theorem 3: for formulas 1 to 5 of the hypersonic aircraft longitudinal motion rigid body model, based on the assumptions 1 to 3, a formula (28) corresponding to a speed subsystem control law and a formula (48) corresponding to an elevator deflection angle control law are adopted, all errors in a closed loop system are finally consistent and bounded, and when the system input exits from saturation, the speed and the altitude tracking error can be ensured to be limited in a preset range, so that the preset transient and steady performance requirements are met. Theorem 3 is demonstrated by the following steps: constructing a Lyapunov function for the whole closed loop system:

Combining equation (30), equation (39), equation (44) and equation (50), deriving equation (58) yields:

wherein , and />All are instruction filter errors;all of them observe errors with a linear extended state observer (linear extended state observer, abbreviated as LESO). The invention designs second-order LESO for disturbance d aiming at disturbance existing in a speed subsystem and a height subsystem respectively ₁ ,d ₂ ,d ₃ The observation and compensation are carried out, specifically:

/>

wherein ,for the estimated value of the aircraft speed V +.>Is the estimated value of track dip angle gamma, +.>Is an estimated value of pitch angle rate q; />As interference term d _i (i=1, 2, 3); l (L) _V1 ,l _V2 ,l _γ1 ,l _γ2 ,l _q1 ,l _q2 All are positive parameters, and the invention adopts a bandwidth configuration method, so that the parameters meet [ l ] _V1 ,l _V2 ]＝[ω _V0 a ₁ ,ω _V0 a ₂ ]，[l _γ1 ,l _γ2 ]＝[ω _γ0 a ₁ ,ω _γ0 a ₂ ]，[l _q1 ,l _q2 ]＝[ω _q0 a ₁ ,ω _q0 a ₂ ]；ω ₀ Representing the bandwidth of the observer, parameter a _i ＝3！/i！·(3-i)！(i＝1,2)；f _V 、g _V 、f _γ 、g _γ 、f _q and g_q As an alternative symbol, reference may be made to the above embodiments for specific formulas; phi represents the fuel equivalent ratio, theta represents the pitch angle, q represents the pitch angle rate, gamma represents the track pitch angle, delta _e Indicating the elevator deflection angle. The present invention makes the following assumption 4 for the demonstration of the convergence of the LESO: LESO observation error- >Is bounded and has an unknown constant +.>So that

Further, according to assumption 3 above, there is an unknown constant N _i (i=1, 2, 3) > 0 such that |η _γ |≤N ₁ ，|η _θ |≤N ₂ ，|η _q |≤N ₃ The method comprises the steps of carrying out a first treatment on the surface of the According to hypothesis 4 above, there is an unknown constantSo that

Note that in formula (59):

in connection with equation (63), equation (59) can be reduced to:

let Λ= [ epsilon ] _V ,e _γ ,e _θ ,υ _q ] ^T ，From the previous analysis, there is a constant N _W > 0, such that->Thus, equation (64) may be further derived:

then whenWhen equation (65) has:

it can thus be stated that W is bounded and epsilon can be derived from the definition of W _V ,e _γ ,e _θ ,υ _q Is bounded. In the invention, when the track dip angle gamma is implemented to gamma _d Conversion error epsilon during tracking of (a) _h (t) satisfyI.e. epsilon _h (t) is bounded. Therefore, according to the above theorem 1, it is known that when γ→γ _d Height tracking error e _h The preset transient and steady state performance requirements are met. From e _γ Bounded availability epsilon _h ∈l _∞ . According to the above theorem 1, from ε _V ,ε _h Is bounded and can obtain v _V ,e _h Is bounded and meets preset transient and steady state performance requirements.

When the system input exits saturation, the first auxiliary system, equation (31), becomesAuxiliary variable xi _V 0, i.e. v _V →e _V Further obtain error e _V The preset transient and steady state performance requirements are met.

Constructing a Lyapunov function for a second auxiliary system, formula (47) construction And deriving to obtain:

when the system inputsAfter exiting saturation, delta _ed ∈[δ _min ,δ _max ]At this timeδ _e ＝H _δ (χ _δ1 ) Therefore, in the formula (67), there is (δ) _e -δ _ed )∈l _∞ . When the auxiliary variable |ζ _q |≥|(δ _e -δ _ed )/k _qξ2 When I, it is clear that equation (67) can be reduced to +.>Thus xi _q Bounded, thereby obtaining e _q Is bounded.

According to the preset performance control method for the hypersonic aircraft, the output tracking error is enabled to have smaller overshoot on the basis of improving the steady state and transient performance of the hypersonic aircraft system by designing a new preset performance function; meanwhile, by constructing the limited instruction filter, the amplitude and the speed of the system input are ensured to meet the limited requirement, so that good tracking performance can be provided on the basis of solving the problem of limited amplitude and speed of the system input.

In an embodiment, the validity of the control scheme provided by the invention is verified, and MATLAB simulation is performed by using the speed subsystem controller and the altitude subsystem controller constructed in the embodiment by taking the hypersonic aircraft longitudinal motion rigid body model formulas (1) to (5) as objects. The aircraft model-related parameters may be selected based on existing parameters.

Specifically, the controller parameters: k (k) _V ＝0.1，λ _V ＝0.01，k _h ＝0.1，k _γ ＝0.8，k _θ ＝2，k _q =2; presetting performance parameters: mu (mu) _V ＝0.2，μ _h ＝0.15，σ _V ＝0.6，σ _h =0.5; auxiliary system parameters: k (k) _Vξ ＝0.8，k _qξ1 ＝10，k _qξ2 =0.02; instruction filteringParameters of wave device: τ _γ ＝τ _θ ＝τ _q ＝0.8，τ _q ＝0.5，ω _γ ＝ω _θ ＝10，ω _q ＝25，ω _δ =90; LESO parameters: bandwidth omega _V0 ＝ω _γ0 ＝ω _q0 =5; the initial value of the system output and state is set as follows: v (V) ₀ ＝7702ft/s，h ₀ ＝85000ft，γ ₀ ＝0rad，θ ₀ ＝0.0264rad，q ₀ =0rad/s; disturbance d ₁ ,d ₂ ,d ₃ The included external interference is set as sin (0.2 t), 0.0002sin (0.2 t), 0.1sin (0.2 t) respectively; consider a system parameter perturbation of +20%.

Control input constraints are set to be phi epsilon 0.05,1.5 respectively]，δ _e ∈[-30°,30°]，Setting the speed and the height step command to be +.>And generates signal command V by the following filters respectively _d ,h _d ：

In order to verify the superiority of the control scheme (denoted as control scheme a) proposed by the present invention, a comparison simulation of an adaptive anti-saturation control scheme (denoted as control scheme B) in the prior art was performed. In order to embody comparative "fairness", the control gain parameter value of control scheme B is the same as control scheme a, and the LESO of the same parameters is employed to observe and compensate for system disturbances.

Fig. 4 is a schematic diagram of comparing speed and tracking error curves provided by the present invention, fig. 5 is a schematic diagram of comparing altitude and tracking error curves provided by the present invention, fig. 6 is a schematic diagram of comparing system state variable curves provided by the present invention, fig. 7 is a schematic diagram of comparing system state variable tracking error curves provided by the present invention, fig. 8 is a schematic diagram of comparing fuel equivalent ratio curves provided by the present invention, fig. 9 is a schematic diagram of comparing elevator deflection angle curves provided by the present invention, fig. 10 is a schematic diagram of comparing elevator deflection angle rate curves provided by the present invention, fig. 11 is a schematic diagram of comparing auxiliary variable curves provided by the present invention, fig. 12 is a schematic diagram of comparing LESO observation curves provided by the present invention, and simulation results can be referred to fig. 4 to 12. Obviously, both control schemes enable the system to achieve stable tracking of instructions (as shown in fig. 4-7). However, referring to fig. 4 and 5, the speed error and altitude error curves of the control scheme a are always within the preset range, and the error convergence speed is better than that of the control scheme B, which indicates that the transient performance of the control system under the control scheme a is better. Referring to fig. 8 to 10, the control input curves of the control scheme a and the control scheme B meet the amplitude constraint condition, but only the control scheme a can meet the rate limiting requirement of the elevator deflection angle, which illustrates the limited instruction filter constructed by the control scheme a, can effectively limit the amplitude and the rate of the control input, and ensure that the actual output of the executing mechanism meets the limiting condition.

Further, referring to fig. 11, when the control input is saturated, the auxiliary variable rapidly responds to compensate for tracking errors, ensuring the stability of the system; when the system exits saturation, the auxiliary variable quickly converges to zero. Referring to fig. 12, the LESO can realize quick and effective observation on system disturbance, which indicates that the system has certain anti-interference capability. In summary, by comparing, the control scheme provided by the invention solves the problem of limited control input amplitude and rate, and simultaneously enables the system to have good transient and steady performance.

Aiming at the hypersonic aircraft tracking performance problem considering input amplitude and rate limitation, the invention provides a preset performance control scheme based on a limited instruction filter. In order to improve transient and steady state performances of the system, a preset performance inversion controller is designed, and the overshoot of tracking errors is smaller by designing a new performance function; secondly, introducing a command filter to treat the problem of difficult derivation in the design of the inversion controller, constructing a limited command filter to restrict the system control law aiming at the problem of limited input, ensuring that the control input meets the limit requirements of amplitude and speed, and carrying out corresponding theoretical evidence; in addition, the uncertainty of system parameters and external interference are considered, and a linear expansion state observer is adopted for observation and compensation. Based on Lyapunov stability theory, it is proved that all tracking errors of the system are finally consistent and bounded. Finally, the effectiveness of the method is verified through simulation.

Fig. 13 is a schematic structural diagram of a preset performance control system for a hypersonic aircraft, as shown in fig. 13, provided by the invention, and including a performance function construction module 1301, a speed subsystem controller construction module 1302, an altitude subsystem controller construction module 1303 and a preset performance control module 1304, where the performance function construction module 1301 is configured to construct a preset performance function of the hypersonic aircraft based on tracking error; the speed subsystem controller construction module 1302 is configured to construct a speed subsystem controller according to the preset performance function and the saturation function; the altitude subsystem controller construction module 1303 is configured to construct an altitude subsystem controller according to the preset performance function and the limited instruction filter by an inversion control method; the preset performance control module 1304 is configured to obtain a state initial value of the hypersonic aircraft at a current moment, and perform tracking control on the hypersonic aircraft according to the speed subsystem controller and the altitude subsystem controller; the saturation function is constructed according to the fuel oil equivalent ratio and the deflection angle of the elevator; the limited instruction filter is constructed based on the input saturation problem and according to the ideal control input value of the elevator deflection angle.

According to the preset performance control system for the hypersonic aircraft, provided by the invention, the output tracking error has smaller overshoot on the basis of improving the steady state and transient performance of the hypersonic aircraft system by designing a new preset performance function; meanwhile, by constructing the limited instruction filter, the amplitude and the speed of the system input are ensured to meet the limited requirement, so that good tracking performance can be provided on the basis of solving the problem of limited amplitude and speed of the system input.

The system provided in the embodiment of the present invention is used for executing the above method embodiments, and specific flow and details refer to the above embodiments, which are not repeated herein.

Fig. 14 is a schematic structural diagram of an electronic device according to the present invention, as shown in fig. 14, the electronic device may include: a processor 1401, a communication interface 1402, a memory 1403 and a communication bus 1404, wherein the processor 1401, the communication interface 1402 and the memory 1403 perform communication with each other through the communication bus 1404. The processor 1401 may invoke logic instructions in the memory 1403 to execute a preset performance control method for a hypersonic aircraft, the method comprising: based on tracking errors, constructing a preset performance function of the hypersonic aircraft; constructing a speed subsystem controller according to the preset performance function and the saturation function; constructing a height subsystem controller through an inversion control method according to the preset performance function and the limited instruction filter; acquiring a state initial value of the hypersonic aircraft at the current moment, and tracking and controlling the hypersonic aircraft according to the speed subsystem controller and the altitude subsystem controller; the saturation function is constructed according to the fuel oil equivalent ratio and the deflection angle of the elevator; the limited instruction filter is constructed based on the input saturation problem and according to the ideal control input value of the elevator deflection angle.

Further, the logic instructions in the memory 1403 described above may be implemented in the form of software functional units and may be stored in a computer readable storage medium when sold or used as a stand alone product. Based on this understanding, the technical solution of the present invention may be embodied essentially or in a part contributing to the prior art or in a part of the technical solution, in the form of a software product stored in a storage medium, comprising several instructions for causing a computer device (which may be a personal computer, a server, a network device, etc.) to perform all or part of the steps of the method according to the embodiments of the present invention. And the aforementioned storage medium includes: a usb disk, a removable hard disk, a Read-only memory (ROM), a random access memory (RAM, randomAccessMemory), a magnetic disk, or an optical disk, or other various media capable of storing program codes.

In another aspect, the present invention also provides a computer program product comprising a computer program stored on a non-transitory computer readable storage medium, the computer program comprising program instructions which, when executed by a computer, are capable of performing the preset performance control method for hypersonic aircraft provided by the methods described above, the method comprising: based on tracking errors, constructing a preset performance function of the hypersonic aircraft; constructing a speed subsystem controller according to the preset performance function and the saturation function; constructing a height subsystem controller through an inversion control method according to the preset performance function and the limited instruction filter; acquiring a state initial value of the hypersonic aircraft at the current moment, and tracking and controlling the hypersonic aircraft according to the speed subsystem controller and the altitude subsystem controller; the saturation function is constructed according to the fuel oil equivalent ratio and the deflection angle of the elevator; the limited instruction filter is constructed based on the input saturation problem and according to the ideal control input value of the elevator deflection angle.

In yet another aspect, the present invention further provides a non-transitory computer readable storage medium having stored thereon a computer program which, when executed by a processor, is implemented to perform the preset performance control method for a hypersonic aircraft provided by the above embodiments, the method comprising: based on tracking errors, constructing a preset performance function of the hypersonic aircraft; constructing a speed subsystem controller according to the preset performance function and the saturation function; constructing a height subsystem controller through an inversion control method according to the preset performance function and the limited instruction filter; acquiring a state initial value of the hypersonic aircraft at the current moment, and tracking and controlling the hypersonic aircraft according to the speed subsystem controller and the altitude subsystem controller; the saturation function is constructed according to the fuel oil equivalent ratio and the deflection angle of the elevator; the limited instruction filter is constructed based on the input saturation problem and according to the ideal control input value of the elevator deflection angle.

The apparatus embodiments described above are merely illustrative, wherein the elements illustrated as separate elements may or may not be physically separate, and the elements shown as elements may or may not be physical elements, may be located in one place, or may be distributed over a plurality of network elements. Some or all of the modules may be selected according to actual needs to achieve the purpose of the solution of this embodiment. Those of ordinary skill in the art will understand and implement the present invention without undue burden.

From the above description of the embodiments, it will be apparent to those skilled in the art that the embodiments may be implemented by means of software plus necessary general hardware platforms, or of course may be implemented by means of hardware. Based on this understanding, the foregoing technical solution may be embodied essentially or in a part contributing to the prior art in the form of a software product, which may be stored in a computer readable storage medium, such as ROM/RAM, a magnetic disk, an optical disk, etc., including several instructions for causing a computer device (which may be a personal computer, a server, or a network device, etc.) to execute the method described in the respective embodiments or some parts of the embodiments.

Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present invention, and are not limiting; although the invention has been described in detail with reference to the foregoing embodiments, it will be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A preset performance control method for a hypersonic aircraft, comprising:

the saturation function is constructed according to the fuel oil equivalent ratio and the deflection angle of the elevator; the limited instruction filter is constructed based on the input saturation problem according to an ideal control input value of the deflection angle of the elevator;

according to the speed subsystem model, deriving a speed error transformation function, and constructing a speed subsystem control law according to the derived speed error transformation function:

the fuel equivalence ratio saturation function is:

wherein the constant phi _max and Φ_min The upper limit and the lower limit of the amplitude value of the fuel oil equivalent ratio phi are respectively set;

wherein ,q_d A pitch angle rate command;

the elevator deflection angle saturation function is as follows:

Constructing an elevator deflection angle control law:

wherein ,k_q > 0, a relevant design parameter for pitch angle rate;for the third interference term d ₃ Is χ _q2 Derivative of the third virtual control law +.>Estimate of k _qξ1 and k_qξ2 Is an auxiliary system parameter;

2. The method for pre-set performance control of a hypersonic aircraft according to claim 1 wherein the constructing a pre-set performance function of a hypersonic aircraft based on tracking errors comprises:

p ₂ (t)＜e(t)＜p ₁ (t)；

wherein ,p₁(t) and p₂ (t) represents a preset performance function, e (0) represents a tracking error at an initial time,representing an existing performance function; />And mu is constant, < >>Is a steady state value +.>And- >

3. The method for pre-set performance control of a hypersonic aircraft according to claim 2 wherein the constructing a speed subsystem controller from the pre-set performance function and saturation function comprises:

υ _V ＝e _V -ξ _V ；

e _V ＝V-V _d ；

p _V2 ＜υ _V ＜p _V1 ；

wherein ,ε_V Representing speed change errorsDifference;

wherein Φ represents the fuel equivalence ratio, d ₁ The first disturbance term is represented by a, the attack angle is represented by D, the resistance is represented by g, the gravitational acceleration is represented by gamma, the track inclination angle is represented by m, the mass is represented by T ₀(α) and T_Φ And (α) represents a thrust-related aerodynamic parameter.

4. A preset performance control method for a hypersonic aircraft according to claim 3 wherein before the constructing a speed subsystem controller from the speed subsystem control law based on the fuel equivalence ratio saturation function, the method further comprises:

where Φ represents the actual input value of the fuel equivalence ratio.

5. The preset performance control method for a hypersonic aircraft according to claim 2 wherein the constructing an altitude subsystem controller by an inversion control method according to the preset performance function and a constrained-instruction filter comprises:

e _h ＝h-h _d ；

p _h2 ＜e _h ＜p _h1 ；

wherein ,ε_h Representing a height conversion error;

wherein V represents the speed of the aircraft, gamma represents the track inclination angle, theta represents the pitch angle, q represents the pitch angle speed, and d ₂ Representing a second interference term, d ₃ Representing a third interference term, delta _e Represents the deflection angle of the elevator, L ₀ and L_α Representing lift-related aerodynamic parameters, g representing gravitational acceleration, M representing mass, M _T 、M ₀(α) and Representing the relevant parameters of pitching moment, I _yy Representing moment of inertia.

6. The method for controlling preset performance of a hypersonic vehicle according to claim 1, wherein before the acquiring the state initial value of the hypersonic vehicle at the current moment and performing tracking control on the hypersonic vehicle according to the speed subsystem controller and the altitude subsystem controller, the method further comprises:

wherein ,for the estimated value of the aircraft speed V +.>Is the estimated value of track dip angle gamma, +.>Is an estimated value of pitch angle rate q; />As interference term d _i I=1, 2,3; l (L) _V1 ,l _V2 ,l _γ1 ,l _γ2 ,l _q1 ,l _q2 Are all positive parameters omega ₀ Representing the bandwidth of the observer, parameter a _i =3-! I-! (3-i) ≡! I=1, 2; phi represents the fuel equivalent ratio, theta represents the pitch angle, q represents the pitch angle rate, gamma represents the track pitch angle, delta _e Indicating the elevator deflection angle.

7. A preset performance control system for a hypersonic aircraft, characterized in that the system performs the preset performance control method for a hypersonic aircraft according to any one of claims 1 to 6.

8. An electronic device comprising a memory, a processor and a computer program stored on the memory and executable on the processor, characterized in that the processor, when executing the computer program, carries out the steps of the preset performance control method for a hypersonic aircraft according to any one of claims 1 to 6.

9. A non-transitory computer readable storage medium, on which a computer program is stored, characterized in that the computer program, when being executed by a processor, implements the steps of the preset performance control method for a hypersonic aircraft according to any one of claims 1 to 6.