CN111061283B - Air suction hypersonic aircraft height control method based on feature model - Google Patents
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Abstract
A height control method of an air suction type hypersonic aircraft based on a feature model belongs to the field of air suction type hypersonic aircraft flight control. The method comprises the following steps: firstly, feature modeling is carried out on a longitudinal plane, and then, self-adaptive control law design is carried out based on the obtained feature model. Compared with the prior art, the method provided by the invention is simple and effective, can cope with uncertainty in various forms, and is also suitable for the condition of large maneuvering flight.
Description
Technical Field
The invention belongs to the field of flight control of air suction hypersonic aircrafts, and relates to an air suction hypersonic aircraft height control method based on a feature model.
Background
The dynamic model of the suction hypersonic aircraft has the characteristics of strong coupling, strong nonlinearity, large uncertainty and the like. In order to solve the problem of large uncertainty, the existing flight control method has more self-adaptive control methods. However, these adaptive control methods assume that the uncertainty term of the system is only a few constant or slowly varying parameters, and is in the form of linearization (i.e., linear parameterization uncertainty). This assumption greatly limits the applicability of these methods because the uncertainty forms in real systems are diverse, with both parameterized and non-parameterized uncertainty, and parameterized uncertainty is further divided into linear and non-linear parameterized uncertainty. Therefore, these adaptive control methods are only a subset of the uncertainties. In addition, the existing self-adaptive controller for the characteristic model of the suction hypersonic aircraft is only suitable for the condition of stable flight.
Disclosure of Invention
The invention solves the technical problems that: the method comprises the steps of firstly carrying out feature modeling on a longitudinal plane, and then carrying out self-adaptive control law design based on the obtained feature model. Compared with the prior art, the method provided by the invention is simple and effective, can cope with uncertainty in various forms, and is also suitable for the condition of large maneuvering flight.
The technical scheme of the invention is as follows:
a method for controlling the altitude of an air suction hypersonic aircraft based on a feature model comprises the following steps:
s1, setting a flight envelope of an aircraft and a boundary of uncertain parameters;
s2, selecting a height as output, selecting an attack angle, an elevator deflection angle and a duck wing deflection angle as input on the basis of an outer ring subsystem formed by the height and the track angle, and establishing a second-order characteristic model of the outer ring subsystem; determining the boundary of the characteristic parameters of the second-order characteristic model of the outer ring subsystem according to the flight envelope and the boundary of the uncertain parameters;
s3, selecting a pitch angle as output, selecting an elevator deflection angle, a duck wing deflection angle and a gas ratio as input, and establishing a second-order characteristic model of the inner ring subsystem based on the pitch angle and the pitch angle rate; determining the boundary of the characteristic parameters of the second-order characteristic model of the inner ring subsystem according to the flight envelope and the boundary of the uncertain parameters;
s4, according to a second-order feature model of the outer ring subsystem, selecting an attack angle and a duck wing deflection angle as control inputs, and obtaining an attack angle instruction and a duck wing deflection angle control law for height tracking;
s5, obtaining a pitch angle instruction by utilizing the track angle and the pitch angle instruction according to the relation between the longitudinal plane pitch angle and the pitch angle;
s6, according to a second-order characteristic model of the inner ring subsystem, selecting an elevator deflection angle as a control input to obtain an elevator deflection angle control law, wherein the elevator deflection angle control law is used for tracking the pitch angle instruction in S5; and the altitude control of the aircraft is completed by utilizing the duck wing deflection angle control law and the elevator deflection angle control law.
Preferably, the boundary of the characteristic parameters of the second-order characteristic model of the outer ring subsystem in the step S2 and the boundary of the characteristic parameters of the second-order characteristic model of the inner ring subsystem in the step S3 are calculated by using, but not limited to, a monte carlo targeting method.
Preferably, in S2, the feature parameters in the second-order feature model of the outer ring subsystem are identified by using a projection gradient identification algorithm or a least squares identification algorithm.
Preferably, in S2, the second-order feature model of the outer ring subsystem is:
h(k+1)＝f 1h (k)h(k)+f 2h (k)h(k-1)+g 1h (k)u 1 (k)+g 2h (k)u 2 (k)+g 3h (k)u 3 (k)
where k corresponds to the kth sampling period, u 1 、u 2 、u 3 Is a control input, f 1h (k)、f 2h (k)、g 1h (k)、g 2h (k) And g 3h (k) All are time-varying characteristic parameters, and the system output of the outer ring subsystem is h (k).
Preferably, in S3, the second-order feature model of the inner ring subsystem is:
θ(k+1)＝2θ(k)-θ(k-1)+g 1θ (k)u 2 (k)+g 2θ (k)u 3 (k)+g 3θ (k)u 4 (k)+σ θ (k)
wherein the method comprises the steps of
Where k corresponds to the kth sampling period, g 1θ 、g 2θ And g 3θ As a characteristic parameter, u 2 、u 3 Is a control input, the system output of the inner loop subsystem is (k), sigma θ In order for the interference term to be a term of interference,represents dynamic pressure, and Φ represents fuel equivalent ratio.
Preferably, S5, the relationship between the angle of attack of the longitudinal plane and the pitch angle is:
α＝θ-γ
where α represents the aircraft angle of attack, θ represents the pitch angle, and γ represents the track angle.
Compared with the prior art, the invention has the beneficial effects that:
(1) Compared with the traditional self-adaptive control method, the method does not need to be uncertain into a linear parameterized form, is also suitable for the condition of large maneuvering flight, and has wider application range.
(2) The controller designed by the traditional adaptive control method is usually in a continuous time form, and is usually a discrete-time sampling controller in engineering practice, so that discretization is also needed. Compared with the traditional self-adaptive control method, the controller obtained by the method is a sampling controller and can be directly used for engineering application.
(3) Compared with the traditional self-adaptive control method, the controller designed by the method has the advantages of simple structure and high reliability.
Drawings
FIG. 1 is a flow chart of the method of embodiment 1 of the present invention;
fig. 2 is a flow chart of the method of embodiment 2 of the present invention.
Detailed Description
Example 1:
the method for controlling the altitude of the air suction hypersonic aircraft based on the feature model is shown in fig. 1, and comprises the following steps:
s1, setting a flight envelope of an aircraft and a boundary of uncertain parameters;
s2, selecting a height as output, selecting an attack angle, an elevator deflection angle and a duck wing deflection angle as input on the basis of an outer ring subsystem formed by the height and the track angle, and establishing a second-order characteristic model of the outer ring subsystem; determining the boundary of the characteristic parameters of the second-order characteristic model of the outer ring subsystem according to the flight envelope and the boundary of the uncertain parameters;
s3, selecting a pitch angle as output, selecting an elevator deflection angle, a duck wing deflection angle and a gas ratio as input, and establishing a second-order characteristic model of the inner ring subsystem based on the pitch angle and the pitch angle rate; determining the boundary of the characteristic parameters of the second-order characteristic model of the inner ring subsystem according to the flight envelope and the boundary of the uncertain parameters;
s4, according to a second-order feature model of the outer ring subsystem, selecting an attack angle and a duck wing deflection angle as control inputs, and obtaining an attack angle instruction and a duck wing deflection angle control law for height tracking;
s5, obtaining a pitch angle instruction by utilizing the track angle and the pitch angle instruction according to the relation between the longitudinal plane pitch angle and the pitch angle;
s6, according to a second-order characteristic model of the inner ring subsystem, selecting an elevator deflection angle as a control input to obtain an elevator deflection angle control law, wherein the elevator deflection angle control law is used for tracking the pitch angle instruction in S5; and the altitude control of the aircraft is completed by utilizing the duck wing deflection angle control law and the elevator deflection angle control law.
Calculating the boundary of the characteristic parameters of the second-order characteristic model of the outer ring subsystem in the step S2 and the boundary of the characteristic parameters of the second-order characteristic model of the inner ring subsystem in the step S3 by using a Monte Carlo targeting method but not limited to the method.
And S2, identifying the characteristic parameters in the second-order characteristic model of the outer ring subsystem by using a projection gradient identification algorithm or a least square identification algorithm. The second-order characteristic model of the outer ring subsystem is as follows:
h(k+1)＝f 1h (k)h(k)+f 2h (k)h(k-1)+g 1h (k)u 1 (k)+g 2h (k)u 2 (k)+g 3h (k)u 3 (k)
where k corresponds to the kth sampling period, u 1 、u 2 、u 3 Is a control input, f 1h (k)、f 2h (k)、g 1h (k)、g 2h (k) And g 3h (k) All are time-varying characteristic parameters, and the system output of the outer ring subsystem is h (k).
S3, the second-order characteristic model of the inner ring subsystem is as follows:
θ(k+1)＝2θ(k)-θ(k-1)+g 1θ (k)u 2 (k)+g 2θ (k)u 3 (k)+g 3θ (k)u 4 (k)+σ θ (k)
wherein the method comprises the steps of
Where k corresponds to the kth sampling period, g 1θ 、g 2θ And g 3θ As a characteristic parameter, u 2 、u 3 Is a control input, the system output of the inner loop subsystem is (k), sigma θ In order for the interference term to be a term of interference,represents dynamic pressure, and Φ represents fuel equivalent ratio.
S5, the relation between the attack angle and the pitch angle of the longitudinal plane is as follows:
α＝θ-γ
where α represents the aircraft angle of attack, θ represents the pitch angle, and γ represents the track angle.
Example 2:
as shown in fig. 2, the height control design process of the present embodiment is: firstly, carrying out feature modeling on a guide outer ring 'height-track angle' subsystem, then carrying out feature modeling on a gesture inner ring 'pitch angle-pitch angle rate' subsystem, on the basis, designing a self-adaptive control law based on a feature model aiming at the guide outer ring to obtain a pitch angle instruction and a duck wing deflection angle control law, and finally, designing a self-adaptive control law based on the feature model aiming at the gesture inner ring to obtain an elevator deflection angle control law so as to track the pitch angle instruction given by the guide outer ring.
Specifically, the longitudinal plane dynamics model of the air suction hypersonic aircraft is as follows:
wherein: h. v, gamma, theta and Q respectively represent altitude, speed, track angle, pitch angle and pitch angle speed; α represents the aircraft angle of attack, α=θ - γ; m, g and I yy Respectively representing the mass, the gravitational acceleration and the moment of inertia of the aircraft; t, L and M yy The engine thrust, lift and pitching moment are respectively expressed as follows:
wherein:represents dynamic pressure, which is associated with V, h; s, z T And->The reference area of the aircraft, the arm of force of the thrust of the engine and the average aerodynamic chord length are respectively represented; phi, delta e And delta c Respectively representing fuel equivalence ratio, elevator deflection angle and duck wing deflection angle, which are control inputs; c in the formula T 、C L And C M The thrust coefficient, the lift coefficient and the pitching moment coefficient of the engine are continuous functions of the respective variables.
The suction hypersonic aircraft mainly flies in the cruising stage, so that the problem of height control in the cruising stage is considered, namely, the elevator deflection angle and the duck wing deflection angle are designed, and the height control is realized. According to the control concept of the inner-outer ring, the formula (1) -formula (4) can be further decomposed into a guidance outer ring of the formula (1) -formula (2) and an attitude inner ring of the formula (3) -formula (4).
A method for controlling the altitude of an air suction hypersonic aircraft based on a feature model comprises the following two parts: feature modeling and adaptive controller design.
For feature modeling, the steps are as follows:
(I) For the outer guidance ring, the selection system input is alpha, delta e And delta c The system output is h. Although C in (6) L (α,δ e ,δ c ) Is a complex expression, but for the cruise phase, its linear term plays a major role, and therefore, there is
In DeltaC L (α,δ e ,δ c ) As the higher-order terms of the argument,respectively lift aerodynamic coefficients. At this time, the relative order of the system is 2, and the following second-order characteristic model is established according to the characteristic model self-adaptive control theory:
h(k+1)＝f 1h (k)h(k)+f 2h (k)h(k-1)+g 1h (k)u 1 (k)+g 2h (k)u 2 (k)+g 3h (k)u 3 (k)(9)
wherein: k corresponds to the kth sampling period, u 1 、u 2 、u 3 Is a control input, expressed as follows
f 1h (k)、f 2h (k)、g 1h (k)、g 2h (k) And g 3h (k) For time-varying characteristic parameters, the expression is as follows
Wherein the method comprises the steps of
T s For sampling time, hereThe derivative representing the speed reference curve, i.e. decoupling the altitude control from the speed control, considers that the speed has tracked the speed reference curve.
(II) in the formula (11) of the step (I), the boundary of the characteristic parameter is related to the boundary of the uncertain parameter and the boundary of the state. In practical application, the boundary of the characteristic parameters can be determined by a Monte Carlo targeting method. Specifically, given a sampling time T s The boundary of uncertain parameters, the boundary of system states and the number of target shooting times, randomly generating uncertain parameters and states in the respective range each time, then calculating the values of the characteristic parameters, and obtaining the boundary of the characteristic parameters with the maximum and minimum values from all target shooting results;
(III) for the gesture inner loop, the system input is selected as delta e 、δ c And phi, the system output is theta. Cruise section, C in formula (7) M The expression is as follows
In DeltaC M (α,δ e ,δ c ) In the case of a higher-order term,are all pitch moment aerodynamic coefficients. Furthermore, C in formula (5) T The expression (alpha, phi) is as follows
In the middle ofThe thrust coefficients corresponding to the linear terms of the fuel gas ratio are adopted; /> The three-order term coefficient of the attack angle, the quadratic term coefficient of the attack angle, the first-order term coefficient of the attack angle and the constant term coefficient of the attack angle are respectively;
at this time, the relative order of the system is 2. According to the self-adaptive control theory of the feature model, firstly, the following second-order feature model is established
θ(k+1)＝2θ(k)-θ(k-1)+g 1θ (k)u 2 (k)+g 2θ (k)u 3 (k)+g 3θ (k)u 4 (k)+σ θ (k) (14)
Wherein u is 2 And u 3 See (10), u 4 The expression of (2) is as follows
g 1θ 、g 2θ And g 3θ As characteristic parameters, the expression is as follows
The sigma theta expression is as follows
The term (14) is called a band interference term "sigma θ "feature model. The presence of the interference term may have an adverse effect on the identification of the characteristic parameter. For this purpose, an output translation transformation is introduced, compressing the disturbances into the characteristic parameters.
Definition of variables
ζ(k)＝θ(k)+χ (18)
Wherein: chi is a normal number, meetsHere, theθ、/>Respectively represent the upper and lower bounds of θ, ε is a positive constant.
Substituting the formula (18) into the formula (14) to obtain
ζ(k+1)＝f 1ζ (k)ζ(k)+f 2ζ (k)ζ(k-1)+g 1ζ (k)u 2 (k)+g 2ζ (k)u 3 (k)+g 3ζ (k)u 4 (k) (1.19)
In the middle of
g 1ζ (k)＝g 1θ (k)
g 2ζ (k)＝g 2θ (k)
g 3ζ (k)＝g 3θ (k)
Equation (1.19) is a feature model corresponding to the inner ring of the gesture.
(IV) in the formula (1.19) of the step (III), the boundary of the characteristic parameter is related to the boundary of the uncertain parameter and the boundary of the state. Obtaining the boundary of the characteristic parameters by using the method provided in the step (II).
For the self-adaptive controller design, the steps are as follows:
(I) For the guidance outer ring, mainly the angle of attack command alpha is designed cmd Height tracking is achieved while the duckwing is adapted to counteract the adverse effects of the elevator. Firstly, a projection gradient identification algorithm or a least square identification algorithm is utilized to identify characteristic parameters f in a characteristic model (9) 1h ,f 2h ,g 1h ,g 2h ,g 3h Obtaining the estimated valueThen, in combination with (10), the controller is designed
In the formula e 1 (k)＝h(k)-h r (k) H is a high tracking error r Is a height reference curve; l (L) 1 ＝0.382，l 2 =0.618 is the golden section coefficient; lambda (lambda) i To control parameters, require andlike numbers, i=1, 2.
(II) further, designing a pitch angle command to be θ according to the relation α=θ - γ between the angle of attack and the pitch angle cmd ＝α cmd +γ。
(III) for the attitude inner ring, the main idea is to design the elevator deflection angle delta e With pitch angle command theta given by tracking guidance outer ring cmd . Firstly, a projection gradient identification algorithm or a least square identification algorithm is utilized to identify a characteristic parameter f in a characteristic model (1.19) 1ζ (k),f 2ζ (k),g 1ζ (k),g 2ζ (k),g 3ζ (k) Obtaining an estimated valueThen, the controller is designed
In the formula e 2 (k)＝ζ(k)-ζ r (k)＝θ(k)-θ cmd (k)；λ 3 To control parameters andsame number.
(IV) combining (1.20), (1.21) and (10) to obtain the elevator deflection angle delta e And a duck wing deflection angle delta c Is an expression of (2). And the altitude control of the aircraft is completed by utilizing the duck wing deflection angle control law and the elevator deflection angle control law.
Claims (6)
1. The method for controlling the height of the air suction hypersonic aircraft based on the feature model is characterized by comprising the following steps of:
s1, setting a flight envelope of an aircraft and a boundary of uncertain parameters;
s2, selecting a height as output, selecting an attack angle, an elevator deflection angle and a duck wing deflection angle as input on the basis of an outer ring subsystem formed by the height and the track angle, and establishing a second-order characteristic model of the outer ring subsystem; determining the boundary of the characteristic parameters of the second-order characteristic model of the outer ring subsystem according to the flight envelope and the boundary of the uncertain parameters;
s3, selecting a pitch angle as output, selecting an elevator deflection angle, a duck wing deflection angle and a gas ratio as input, and establishing a second-order characteristic model of the inner ring subsystem based on the pitch angle and the pitch angle rate; determining the boundary of the characteristic parameters of the second-order characteristic model of the inner ring subsystem according to the flight envelope and the boundary of the uncertain parameters;
s4, according to a second-order feature model of the outer ring subsystem, selecting an attack angle and a duck wing deflection angle as control inputs, and obtaining an attack angle instruction and a duck wing deflection angle control law for height tracking;
s5, obtaining a pitch angle instruction by utilizing the track angle and the pitch angle instruction according to the relation between the longitudinal plane pitch angle and the pitch angle;
s6, according to a second-order characteristic model of the inner ring subsystem, selecting an elevator deflection angle as a control input to obtain an elevator deflection angle control law, wherein the elevator deflection angle control law is used for tracking the pitch angle instruction in S5; and the altitude control of the aircraft is completed by utilizing the duck wing deflection angle control law and the elevator deflection angle control law.
2. The method for controlling the altitude of an air suction hypersonic aircraft based on a characteristic model as claimed in claim 1, wherein: calculating the boundary of the characteristic parameters of the second-order characteristic model of the outer ring subsystem in the step S2 and the boundary of the characteristic parameters of the second-order characteristic model of the inner ring subsystem in the step S3 by using a Monte Carlo targeting method but not limited to the method.
3. The method for controlling the altitude of an air suction hypersonic aircraft based on a characteristic model as claimed in claim 1, wherein: and S2, identifying the characteristic parameters in the second-order characteristic model of the outer ring subsystem by using a projection gradient identification algorithm or a least square identification algorithm.
4. The method for controlling the altitude of an air suction hypersonic aircraft based on a characteristic model as claimed in claim 1, wherein: in S2, the second-order feature model of the outer ring subsystem is:
h(k+1)＝f 1h (k)h(k)+f 2h (k)h(k-1)+g 1h (k)u 1 (k)+g 2h (k)u 2 (k)+g 3h (k)u 3 (k)
where k corresponds to the kth sampling period, u 1 、u 2 、u 3 Is a control input, f 1h (k)、f 2h (k)、g 1h (k)、g 2h (k) And g 3h (k) All are time-varying characteristic parameters, and the system output of the outer ring subsystem is h (k).
5. The method for controlling the altitude of an air suction hypersonic aircraft based on a characteristic model as claimed in claim 1, wherein: s3, the second-order characteristic model of the inner ring subsystem is as follows:
θ(k+1)＝2θ(k)-θ(k-1)+g 1θ (k)u 2 (k)+g 2θ (k)u 3 (k)+g 3θ (k)u 4 (k)+σ θ (k)
wherein the method comprises the steps of
Where k corresponds to the kth sampling period, g 1θ 、g 2θ And g 3θ As a characteristic parameter, u 2 、u 3 Is a control input, the system output of the inner loop subsystem is (k), sigma θ In order for the interference term to be a term of interference,represents dynamic pressure, and Φ represents fuel equivalent ratio.
6. The method for controlling the altitude of an air suction hypersonic aircraft based on a characteristic model according to any one of claims 1 to 5, wherein: s5, the relation between the attack angle and the pitch angle of the longitudinal plane is as follows:
α＝θ-γ
where α represents the aircraft angle of attack, θ represents the pitch angle, and γ represents the track angle.
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