CN107943097B - Aircraft control method and device and aircraft - Google Patents

Aircraft control method and device and aircraft Download PDF

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CN107943097B
CN107943097B CN201711498560.4A CN201711498560A CN107943097B CN 107943097 B CN107943097 B CN 107943097B CN 201711498560 A CN201711498560 A CN 201711498560A CN 107943097 B CN107943097 B CN 107943097B
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CN107943097A (en
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刘昊
马腾
蔡国飙
刘德元
赵万兵
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Beihang University
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    • G05—CONTROLLING; REGULATING
    • G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
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    • G05D1/10—Simultaneous control of position or course in three dimensions
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Abstract

The invention provides an aircraft control method and device and an aircraft; wherein, the method comprises the following steps: acquiring initial flight state parameters output by a power system of an aircraft; generating an initial control signal according to the initial flight state parameter and a preset reference signal; inputting the initial control signal into a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and a reference signal; generating a compensation control signal according to the error signal; and determining throttle valve and rudder deflection angle control signals of the aircraft according to the initial control signal and the compensation control signal so as to control the flight state of the aircraft. The invention can inhibit the influence of various uncertain factors generated by the aircraft in the flight process and improve the tracking control performance of the controller on the flight state of the aircraft; meanwhile, the linear time-invariant control method is easy to implement and high in practicability.

Description

Aircraft control method and device and aircraft
Technical Field
The invention relates to the technical field of aircrafts, in particular to an aircraft control method and device and an aircraft.
Background
High speed aircraft are widely used in a variety of applications because of their ability to quickly, efficiently and reliably enter adjacent spaces. High speed aircraft are subject to non-linearities and uncertainties including parameter uncertainties, non-linear coupling, unstructured and external disturbances during flight, which can severely affect the tracking performance of closed loop control systems in the aircraft, especially when performing supersonic flight missions. In the existing control mode, uncertain influences are often ignored or roughly estimated, so that the uncertain factor inhibition capability of the aircraft is poor, and further the tracking control performance is poor.
Disclosure of Invention
In view of the above, the present invention provides an aircraft control method and apparatus, and an aircraft, so as to suppress the influence of various uncertain factors generated by the aircraft during a flight process, and improve the tracking control performance of a controller on the flight state of the aircraft.
In a first aspect, an embodiment of the present invention provides a method for controlling an aircraft, where the method is applied to a controller of the aircraft; the method comprises the following steps: acquiring initial flight state parameters output by a power system of an aircraft; wherein the initial flight state parameters comprise flight speed and flight altitude; generating an initial control signal according to the initial flight state parameter and a preset reference signal; inputting the initial control signal into a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and a reference signal; generating a compensation control signal according to the error signal; and determining throttle valve and rudder deflection angle control signals of the aircraft according to the initial control signal and the compensation control signal so as to control the flight state of the aircraft.
With reference to the first aspect, an embodiment of the present invention provides a first possible implementation manner of the first aspect, where the step of generating an initial control signal according to the initial flight state parameter and a preset reference signal includes: calculating an initial control signal u by the following formulai,H：
ui,H＝Ki,Hei,i＝1,2,；
Wherein, i-1 represents the flying speed; i-2 represents the flying height; ki,HSuboptimal state feedback gain; e.g. of the typeiIs an error signal.
With reference to the first possible implementation manner of the first aspect, an embodiment of the present invention provides a second possible implementation manner of the first aspect, wherein the suboptimal state feedback gain K is obtained by applying a first bias voltage to the suboptimal state feedback gain Ki,HObtained by the following formula:
Figure BDA0001533404850000021
wherein,
Figure BDA0001533404850000022
ωnis the natural angular frequency;
Figure BDA0001533404850000023
ρ is the atmospheric density; v. oftrimThe speed of the aircraft in the balance state in the cruising flight stage; s is a reference area;
Figure BDA0001533404850000024
is the average aerodynamic chord length; cMeIs the aerodynamic coefficient; i isyyIs the moment of inertia; piIs a symmetric positive solution of the following equation:
Figure BDA0001533404850000031
Figure BDA0001533404850000032
Figure BDA0001533404850000033
a21＝vtrim；a22＝Ttrim/mN/Vtrim；CTβ0and CTβ2Is the aerodynamic coefficient; m is the aircraft mass; zetanIs the damping ratio; t istrimThrust of the aircraft in a balanced state in a cruising flight stage;
Figure BDA0001533404850000034
Figure BDA0001533404850000037
and thetaiIs a weight parameter; qi(i ═ 1,2) is a symmetric positive definite matrix, φi(i ═ 1,2) is the set attenuation factor; the superscript N indicates that the parameter is a nominal parameter.
With reference to the second possible implementation manner of the first aspect, the embodiment of the present invention provides a third possible implementation manner of the first aspectIn one embodiment, the step of generating the compensation control signal according to the error signal includes: calculating a compensation control signal u by the following formulai,R(s)：
Figure BDA0001533404850000035
Wherein s is a Laplace operator; f1(s)＝f1 3/(s+f1)3，
Figure BDA0001533404850000036
f1And f2Is a set filtering parameter which is a normal number; y is2＝h-rh，y1＝v-rv(ii) a v is the speed of the aircraft; h is the height of the aircraft; r isvThe reference speed of the aircraft in the cruising flight stage; r ishThe reference height of the aircraft in the cruising flight stage; gi(s) is from ui,R(s) to yiThe transfer function of (2).
With reference to the third possible implementation manner of the first aspect, the embodiment of the present invention provides a fourth possible implementation manner of the first aspect, where the transfer function G is described abovei(s) obtained by the following formula:
Gi(s)＝Ci(sIi-Ai,H)-1Bi(i＝1,2)；
wherein, Ii(i ═ 1,2) is an identity matrix;
Figure BDA0001533404850000041
Ai,H＝Ai+BiKi,H。
with reference to the fourth possible implementation manner of the first aspect, the embodiment of the present invention provides a fifth possible implementation manner of the first aspect, where the step of determining the throttle valve and rudder deflection angle control signals of the aircraft according to the initial control signal and the compensation control signal to control the flight state of the aircraft includes: calculating a final control signal u for the aircrafti＝ui,H+ui,R(i ═ 1, 2); and controlling the flying speed and/or flying height of the aircraft according to the final control signal.
In a second aspect, an embodiment of the present invention provides a control device for an aircraft, where the device is disposed in a controller of the aircraft; the device comprises: the signal acquisition module is used for acquiring initial flight state parameters output by a power system of the aircraft; wherein the initial flight state parameters comprise flight speed and flight altitude; the first signal generation module is used for generating an initial control signal according to the initial flight state parameter and a preset reference signal; the error signal calculation module is used for inputting the initial control signal to a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and a reference signal; the second signal generation module is used for generating a compensation control signal according to the error signal; and the control module is used for determining a throttle valve and a rudder deflection angle control signal of the aircraft according to the initial control signal and the compensation control signal so as to control the flight state of the aircraft.
With reference to the second aspect, an embodiment of the present invention provides a first possible implementation manner of the second aspect, where the first signal generating module is further configured to: calculating an initial control signal u by the following formulai,H：ui,H＝Ki,HeiI is 1, 2; wherein, i-1 represents the flying speed; i-2 represents the flying height; ki,HFeedback gain for suboptimal state; e.g. of the typeiIs an error signal.
With reference to the first possible implementation manner of the second aspect, an embodiment of the present invention provides a second possible implementation manner of the second aspect, where the first signal generating module is further configured to: suboptimal state feedback gain Ki,HObtained by the following formula:
Figure BDA0001533404850000051
wherein,
Figure BDA0001533404850000052
ωnis the natural angular frequency;
Figure BDA0001533404850000053
ρ is the atmospheric density; v. oftrimThe speed of the aircraft in the balance state in the cruising flight stage; s is a reference area;
Figure BDA0001533404850000054
is the average aerodynamic chord length; cMeIs the aerodynamic coefficient; i isyyIs the moment of inertia; piIs a symmetric positive solution of the following equation:
Figure BDA0001533404850000055
Figure BDA0001533404850000056
Figure BDA0001533404850000057
a21＝vtrim；a22＝Ttrim/mN/Vtrim；CTβ0and CTβ2Is the aerodynamic coefficient; m is the aircraft mass; zetanIs the damping ratio; t istrimThrust of the aircraft in a balanced state in a cruising flight stage;
Figure BDA0001533404850000058
Figure BDA0001533404850000059
and thetaiIs a weight parameter; qi(i ═ 1,2) is a symmetric positive definite matrix, φi(i ═ 1,2) is the set attenuation factor; the superscript N indicates that the parameter is a nominal parameter.
In a third aspect, embodiments of the present invention provide an aircraft, where the aircraft includes a processor and a sensor; the control device of the aircraft is arranged in the processor.
The embodiment of the invention has the following beneficial effects:
according to the control method and device for the aircraft and the aircraft provided by the embodiment of the invention, the initial control signal can be generated according to the initial flight state parameter output by the aircraft and the preset reference signal; after the initial control signal is input into a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and a reference signal; then, according to the error signal, a compensation control signal can be generated; then according to the initial control signal and the compensation control signal, a throttle valve and a rudder deflection angle control signal of the aircraft can be determined, and the flight state of the aircraft is controlled; the method can inhibit the influence of various uncertain factors generated by the aircraft in the flight process by compensating the control signal, and improves the tracking control performance of the controller on the flight state of the aircraft.
Furthermore, the method adopts a linear time-invariant control method, is easy to realize and has strong practicability.
Additional features and advantages of the invention will be set forth in the description which follows, and in part will be obvious from the description, or may be learned by the practice of the invention as set forth above.
In order to make the aforementioned and other objects, features and advantages of the present invention comprehensible, preferred embodiments accompanied with figures are described in detail below.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, and it is obvious that the drawings in the following description are some embodiments of the present invention, and other drawings can be obtained by those skilled in the art without creative efforts.
Fig. 1 is a flowchart of a control method for an aircraft according to an embodiment of the present invention;
fig. 2 is a schematic structural diagram of a flight controller according to an embodiment of the present invention;
FIG. 3 is a diagram illustrating a comparison of singular values of velocity channels for various controllers provided by an embodiment of the present invention;
FIG. 4 is a diagram illustrating a comparison of singular values of the altitude channel for various controllers provided by an embodiment of the present invention;
FIG. 5 is a diagram of H according to an embodiment of the present invention∞A tracking response graph of a speed channel and a height channel of the controller to a reference signal;
FIG. 6 is a graph of the tracking response of the velocity channel and the altitude channel of a linear state feedback controller to a reference signal provided by an embodiment of the present invention;
FIG. 7 is a drawing showing a schematic diagram of H according to an embodiment of the present invention∞Response graphs of track angle, attack angle and roll angular velocity of the controller and the linear state feedback controller;
FIG. 8 is a diagram of a numerical simulation process H according to an embodiment of the present invention∞The controller and the linear state feedback controller are used for generating an input value change schematic diagram of an aircraft power system;
FIG. 9 is a drawing showing a schematic diagram of H according to an embodiment of the present invention∞The controller and the linear state feedback controller compare the tracking error of the speed channel and the tracking error of the height channel;
fig. 10 is a schematic structural diagram of a control device of an aircraft according to an embodiment of the present invention.
Detailed Description
To make the objects, technical solutions and advantages of the embodiments of the present invention clearer, 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, but not all embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
In consideration of the problem that the tracking control performance is poor due to poor uncertain factor inhibition capability of the existing aircraft control mode on the aircraft, the embodiment of the invention provides an aircraft control method and device and the aircraft; the technology can be applied to the flight control process of high-speed aircrafts, unmanned planes and the like; the techniques may be implemented in associated software or hardware, as described by way of example below.
Referring to FIG. 1, a flow chart of a method of controlling an aircraft is shown; the method is applied to a controller of an aircraft; the method comprises the following steps:
step S102, collecting initial flight state parameters output by a power system of an aircraft; wherein the initial flight state parameters comprise flight speed and flight altitude;
step S104, generating an initial control signal according to the initial flight state parameter and a preset reference signal;
step S105, inputting the initial control signal to a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and a reference signal; generally, the controller can process the error signal of flying speed and the error signal of flying height respectively, and then track unmanned aerial vehicle's flying speed and flying height respectively.
Specifically, the initial control signal generates a speed channel control law and an altitude channel control law through a controller, and then is input to the aircraft; due to the effects of non-linearity and uncertainty, there can be errors between the actual state of flight of the aircraft and the reference signal; based on this, the above step S105 calculates an error signal between the current flight state parameter and the reference signal to compensate the initial control signal according to the error signal.
Specifically, the initial control signal u may be calculated by the following formulai,H：
ui,H＝Ki,Hei,i＝1,2,；
Wherein, i-1 represents the flying speed; i-2 represents the flying height; ki,HFeedback gain for suboptimal state; e.g. of the typeiIs an error signal.
The suboptimal state feedback gain Ki,HCan pass throughThe following formula is obtained:
Figure BDA0001533404850000081
wherein,
Figure BDA0001533404850000091
ωnis the natural angular frequency;
Figure BDA0001533404850000092
ρ is the atmospheric density; v. oftr i mThe speed of the aircraft in the balance state in the cruising flight stage; s is a reference area;
Figure BDA0001533404850000093
is the average aerodynamic chord length; cMeIs the aerodynamic coefficient; i isyyIs the moment of inertia;
p is aboveiIs a symmetric positive solution of the following equation:
Figure BDA0001533404850000094
Figure BDA0001533404850000095
Figure BDA0001533404850000096
a21＝vtrim；a22＝Ttrim/mN/Vtrim；CTβ0and CTβ2Is the aerodynamic coefficient; m is the aircraft mass; zetanIs the damping ratio; t istrimThrust of the aircraft in a balanced state in a cruising flight stage;
Figure BDA0001533404850000097
and thetaiIs a weight parameter; qi(i ═ 1,2) is a symmetric positive definite matrix, φi(i ═ 1,2) is the set attenuation factor; the superscript N indicates that the parameter is a nominal parameter.
Step S106, generating a compensation control signal according to the error signal;
specifically, the compensation control signal u can be calculated by the following formulai,R(s)：
Figure BDA0001533404850000099
Wherein s is a Laplace operator; f1(s)＝f1 3/(s+f1)3，
Figure BDA0001533404850000101
f1And f2Is a set filtering parameter which is a normal number; y is1＝v-rv，y2＝h-rh(ii) a v is the speed of the aircraft; h is the height of the aircraft; r isvThe reference speed of the aircraft in the cruising flight stage; r ishThe reference height of the aircraft in the cruising flight stage; gi(s) is from ui,R(s) to yiThe transfer function of (2).
The above transfer function Gi(s) obtained by the following formula:
Gi(s)＝Ci(sIi-Ai,H)-1Bi(i＝1,2)；
wherein, Ii(i ═ 1,2) is an identity matrix;
Figure BDA0001533404850000102
Ai,H＝Ai+BiKi,H。
and S108, determining throttle valve and rudder deflection angle control signals of the aircraft according to the initial control signal and the compensation control signal so as to control the flight state of the aircraft.
This step S108 may be specifically implemented by:
step (1) of carrying out a treatment,calculating a final control signal u for the aircrafti＝ui,H+ui,R，(i＝1,2)；
And (2) controlling the flying speed and/or flying height of the aircraft according to the final control signal.
As can be seen from the above, i ═ 1 represents the flight speed; i-2 represents the flying height; thus, when i is 1, the controller may be based on u1Controlling the flight speed of the aircraft; when i is 2, can be according to u2Controlling the flight speed of the aircraft; of course, the controller may also generate u simultaneously1And u2To simultaneously control the speed and altitude of the aircraft.
According to the control method of the aircraft provided by the embodiment of the invention, the initial control signal can be generated according to the initial flight state parameter output by the aircraft and the preset reference signal; after the initial control signal is input into a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and a reference signal; then, according to the error signal, a compensation control signal can be generated; then according to the initial control signal and the compensation control signal, a throttle valve and a rudder deflection angle control signal of the aircraft can be determined, and the flight state of the aircraft is controlled; the method can inhibit the influence of various uncertain factors generated by the aircraft in the flight process by compensating the control signal, and improves the tracking control performance of the controller on the flight state of the aircraft.
Furthermore, the method adopts a linear time-invariant control method, is easy to realize and has strong practicability.
The embodiment of the invention also provides another aircraft control method based on H∞Realizing a theoretical and robust compensation method; the method can inhibit the influence of strong nonlinearity and effective disturbance including strong coupling, unstructured uncertainty, external interference and the like on the designed closed-loop control system, and generate a robust linear control system which does not meet the matching condition.
Specifically, the method is based first on H∞Theoretical design H∞A controller for controlling the operation of the electronic device,to achieve the desired tracking performance; due to H∞The controller cannot suppress the effects of non-linearities and uncertainties on the closed-loop control system over the entire frequency domain, and therefore, the robust compensator is redesigned to suppress the effects of these equivalent disturbances. H∞The controller and the robust compensator are combined to form a linear time-invariant controller, which can ensure that the tracking errors of the speed channel and the altitude channel of the high-speed aircraft with nonlinearity and multiple uncertainties are finally converged into the neighborhood of a given pair near the origin. In addition, the method has robustness of linear time invariance, is easy to realize in practical application, and has strong practicability.
To implement this method, it is first necessary to build a dynamic model of the high-speed aircraft. The longitudinal dynamics of the high-speed aircraft can be described by the following equations:
Figure BDA0001533404850000111
Figure BDA0001533404850000112
Figure BDA0001533404850000113
Figure BDA0001533404850000114
Figure BDA0001533404850000115
the model includes five state variables [ v, γ, h, α, q]Where v represents velocity, γ represents flight path angle, h represents altitude, α represents angle of attack, q represents pitch rate, m represents aircraft mass, μ represents gravitational constant, I represents altitude, andyydenotes moment of inertia, r ═ h + reWherein r iseIs the radius of the earth; di(i ═ V, γ, q, α, h) are external time-varying atmospheric disturbances (note: in practice, external disturbances refer primarily to aircraft dynamicsAdditional forces and moments introduced by the model. So d in the formula (1)hAnd dαNot of practical significance, they are used in stability analysis introduced into closed-loop control systems), T, L, D and MqRespectively expressing thrust, lift, resistance and pitching moment, and satisfying the following formula:
L＝ρv2SCL/2,
T＝ρv2SCT/2,
D＝ρv2SCD/2,
Figure BDA0001533404850000121
wherein the sum of p, S,
Figure BDA0001533404850000122
respectively representing the atmospheric density, the reference area and the average aerodynamic chord length; cL,CT,CD,CMα,CMeRepresenting thrust coefficient, lift coefficient, drag coefficient, angle of attack coefficient, yaw rate coefficient and pitch rate coefficient, respectively, depending on angle of attack α and rudder deflection angleeAnd satisfies the following:
CL＝CLαα+ΔL,
Figure BDA0001533404850000123
CD＝CDα2α2+CDαα+CD0+ΔD,
CMα＝CMα2α2+CMαα+Cα0+ΔMα,
CMe＝CMe(e-α)+Δe,
Figure BDA0001533404850000124
wherein β is the throttle opening, CLα，CTβ0，CTβ1，CTβ2，CDα2，CDα，CD0，CMα2，CMα，Cα0，CMe，CMq2，CMq，Cq0Is the aerodynamic coefficient, is the unmodeled uncertainty; since it is difficult to determine the analytical relationship between the thrust equivalent and moment coefficients and the angle of attack or throttle opening, the curve fitting technique is used to describe the aerodynamic characteristics when designing the controller, so that a mismatch term Δ between the real model and the control-oriented model occursj. The high speed aircraft engine dynamics model may be described by the following second order system:
Figure BDA0001533404850000125
wherein, βtsc，ωnAnd ξnRespectively representing a throttle opening command, a natural angular frequency and a damping ratio; dβRepresenting an external disturbance acting on the engine. The elastic model of the longitudinal model of the high-speed aircraft can be restrained by an additional duck-shaped layout, and acts on the motion d of the high-speed aircraftβCan be seen as a bounded perturbation.
In the embodiment, the reference speed r of the high-speed aircraft in the cruising flight stage can be realized through the robust feedback controllervAnd a reference height rhThe tracking of (2). The speed, the altitude and the thrust are respectively v at the equilibrium state in the cruising flight stagetrim，htrimAnd Ttrim. Defining two control inputs as u1＝βtsc，u2＝e(ii) a Defining the output altitude error and the speed error as: y is1＝v-rv，y2＝h-rh(ii) a Let e1＝[e1,i]3×1，e2＝[e2,i]4×1Wherein e is1,1＝y1，e1,2＝β，
Figure BDA0001533404850000131
e2,1＝y2，e2,2＝γ，e2,3＝α，e2,4Q; thus the high speed aircraft longitudinal model represented by equations (1) - (3) can be rewritten as:
Figure BDA0001533404850000132
yi＝Ciei,i＝1,2,(5)
Figure BDA0001533404850000133
Figure BDA0001533404850000134
Figure BDA0001533404850000135
a21＝vtrim,a22＝Ttrim/mN/Vtrim,
Figure BDA0001533404850000136
wherein the upper-scale N represents a parameter as a nominal parameter; delta1＝[Δ1,i]3×1And Δ2＝[Δ2,i]4×1Are equivalent perturbations including parameter uncertainty, unmodeled uncertainty, nonlinearity, coupling, and external disturbances.
After the establishment of the dynamic model of the high-speed aircraft is completed, an aircraft controller needs to be established so as to realize the control method of the aircraft; specifically, criterion H may be∞Control theory and robust compensation strategy are combined to obtain expected tracking effect and weaken equivalent disturbance deltai(i-1, 2). Referring to fig. 2, a schematic diagram of a flight controller is shown; input u of flight controlleri(i-1, 2) includes the criterion H∞Controlled output value ui,H(i ═ 1,2) and output value u based on robust compensation theoryi,R(i＝1,2)，ui(i ═ 1,2) represents the following:
ui＝ui,H+ui,R,i＝1,2. (6)
first, use H∞Control theory to design H of height and speed channel∞And a controller. Let zi(i ═ 1,2) represents output performance, and has the following form: z is a radical ofi＝Ci,eei+Ci,uui,H,i＝1,2；
The systematic error is:
Figure BDA0001533404850000141
Figure BDA0001533404850000142
Figure BDA0001533404850000143
wherein, c1i(i＝1,2,3)，c2i(i ═ 1,2,3,4) and θi(i ═ 1,2) is a weight parameter; the uncertain error system (i.e. formula (7)) has suboptimal state feedback gain Ki,HK is the same asi,HCan be composed of
Figure BDA0001533404850000144
Is calculated to obtain, wherein Pi(i ═ 1,2) is a symmetric positive solution of the following equation:
Figure BDA0001533404850000145
wherein Q isi(i ═ 1,2) is a symmetric positive definite matrix, φi(i ═ 1,2) is a given attenuation factor;
from above, H∞The state feedback control law can ultimately be given by the following equation:
ui,H＝Ki,Hei,i＝1,2, (8)
wherein, K1,H＝[k1,i]1×3，K2,H＝[k2,i]1×4。
Second, the robust compensator is designed to attenuate the uncertainty Δi(i ═ 1,2) impact on closed loop control systems;
from equations (5) and (6), it can be found that:
Figure BDA0001533404850000151
yi＝Ciei,i＝1,2, (9)
wherein A isi,H＝Ai+BiKi,H(ii) a Let Gi(s) (i ═ 1,2) represents the altitude and velocity path from input ui,RTo the output yiA transfer function of (2), and Gi(s)＝Ci(sIi-Ai,H)-1Bi(I ═ 1,2) where Ii(i ═ 1,2) represents an identity matrix. From equation (9):
yi(s)＝Gi(s)ui,R(s)+Ci(sIi-Ai,H)-1(ei(0)+Δi(s)),i＝1,2. (10)
in order to attenuate the equivalent interference deltai(i ═ 1,2), robustly compensating for input ui,R(i ═ 1,2) can be provided in the form of:
Figure BDA0001533404850000152
as is apparent from the formula (11)
Figure BDA0001533404850000153
Can completely cancel deltaiThe influence of (c). Equation (9) due to Δi(i-1, 2) relates to state derivation, so
Figure BDA0001533404850000154
It cannot be directly implemented, so a low-pass robust filter needs to be introduced to remove the influence of these derivatives, the low-pass filter having the form:
Figure BDA0001533404850000155
F1(s) and F2(s) acting on the velocity and altitude paths, respectively, wherein the filter parameter f1And f2Is the undetermined normal number. Thus robustly compensating the input ui,R(i ═ 1,2) can be redesigned as follows:
Figure BDA0001533404850000156
since Δ cannot be directly measured or obtainedi(i ═ 1,2), the robust compensation controller can obtain the following form by substituting equation (10) into equation (13):
Figure BDA0001533404850000157
in order to prove the tracking performance of the control method of the aircraft, the present embodiment theoretically proves the tracking control performance of the flight controller, and specifically, the flight controller (which may also be referred to as a robust controller) of the present embodiment is explained by using a bode diagram compared with the flight controller using only H∞Controller (also can be called standard H)∞State feedback controller). The following example is used for discussion: calculating two controller speed channel slave delta1,3To y1And altitude passage from2,4To y2The singular values of the transfer function of (2) are analyzed in comparison in a bode plot. Wherein, standard H∞The parameters of the state feedback controller are selected as follows: c. C11＝8×105,c12＝106,c13＝10,c21＝5,c22＝104,c23＝78,c24＝16.2,θ1＝1,θ2＝0.05,φ1＝300,φ2＝3×105。
Referring to FIG. 3 for a graph showing the comparison of singular values of the velocity channel for various controllers and FIG. 4 for a graph showing the comparison of singular values of the height channel for various controllers; as can be seen from fig. 3 and 4, in the low frequency band, the present embodiment providesFlight controller ratio H of supply∞The singular values of the controller are small and, after adding the robust filter shown in equation (12), the singular values of the low frequency band may become smaller if larger filter parameters are selected. This illustrates a robust controller ui,R(i ═ 1,2) can limit the equivalent interference Δ1,3And Δ2,4Impact on control performance. The same method can be implemented in other input-output transfer functions of high-speed aircraft.
From equations (10) - (13), the following closed-loop transfer matrices for the velocity and altitude paths can be derived:
yi(s)＝(1-Fi(s))Ci(sIi-Ai,H)-1Δi(s),i＝1,2,
if the robust filter parameter takes a sufficiently large value, the robust filter becomes an all-pass filter and the gain becomes 1, so that the equivalent interference ΔiThe influence of (i ═ 1,2) can be suppressed in the entire frequency domain.
Because of Δi(i-1, 2) cannot be considered as an external disturbance only, and therefore cannot be assumed to be bounded, and at the same time, the robust filtering parameter f in practical applicationsiAgain, (i ═ 1,2) cannot be made large enough, so the robust stability of control systems designed under non-linearity and uncertainty is discussed further.
Because the high speed aircraft model (i.e., equation (9)) does not satisfy the matching condition, the system variables are transformed as follows before discussing system stability: definition of
x1＝[x1,i]3×1,x2＝[x2,i]4×1
Wherein x is1,1＝e1,1,x2,1＝e2,1,
Figure BDA0001533404850000171
Figure BDA0001533404850000172
And is
Figure BDA0001533404850000173
Further rewritten as follows:
Figure BDA0001533404850000174
yi＝Ciei,i＝1,2. (15)
the norm used in this example is as follows:
Figure BDA0001533404850000175
Figure BDA0001533404850000176
wherein,
Figure BDA0001533404850000177
the equivalent perturbation is assumed to have the following bounded norm:
Figure BDA0001533404850000178
wherein λ isΔx4i,λΔx3i,λΔx2i,λΔx1i,λΔciIs a normal number. The tracking performance of a closed loop control system designed in this way can be described by the following theorem:
theorem 1: under the above assumptions, for any given initial state x (0) with a bounded initial state, any positive constant f can be foundminAnd TminIf f is satisfiedi≥fmin(i ═ 1,2), then any state x is bounded, and
Figure BDA0001533404850000179
|yiless than or equal to (i-1, 2); further if x (0) is 0, then
Figure BDA00015334048500001710
|yi|≤(i＝1,2)。
Theorem 1 is demonstrated as follows:
from equation (15), the following can be derived:
||xi||∞≤λxi(0)+i||Δi||∞,i＝1,2, (17)
wherein λ isxi(0)Is a normal number and satisfies:
Figure BDA0001533404850000181
i＝||(sIi-Ai,H)-1(1-Fi)||1. It can be seen thatxi(0)Depending on the initial conditions, if the initial state x (0) is bounded, there is a positive constant λxi(0)Satisfy the requirement of
Figure BDA0001533404850000182
It was mentioned above that if the filter parameter is chosen to be large positive, the filter Fi(s) (i ═ 1,2) gain close to 1, so thatiMay become smaller. The normal number f can be obtained by referring to articles such as Liuhao and the likemin1And λFor any fi≥fmin1(i＝1,2)，||||∞≤λThe/f is true.
Wherein maxi i(ii) a Let lambdax(0)＝maxiλxi(0)，λΔxj＝maxiλΔxji(j＝1,2,3,4)，λΔc＝maxiλΔci，
Figure BDA0001533404850000183
From equations (16) - (18), the following inequalities can be derived:
Figure BDA0001533404850000184
if f isi(i ═ 1,2) satisfies:
Figure BDA0001533404850000185
fi≥λ,i＝1,2. (20)
substituting equations (17) and (19) into (16) can yield:
Figure BDA0001533404850000186
from equations (17) and (18), it is possible to obtain:
Figure BDA0001533404850000187
wherein λ isxfIs a normal number and satisfies:
Figure BDA0001533404850000188
the first inequality of equation (20) determines the attraction zone as follows:
{x:||x||∞≤ξmax}, (23)
wherein, ξmaxIs the largest positive root of the following equation:
Figure BDA0001533404850000189
from the above expression, there are positive real numbers fmin2And satisfies the following conditions: when f isi≥fmin2(i is 1, 2):
ξmax>λx(0),
ξmax≥||x(0)||∞. (24)
thus, for any fi≥fmin3(i ═ 1,2) satisfies:
Figure BDA0001533404850000191
thus, the inequality (19) holds.
Finally, from (17), (18), and (22) can be obtained:
Figure BDA0001533404850000192
wherein, cik(i ═ 1,2) is a vector of the k-th line 1 and the other lines 0. Order to
Figure BDA0001533404850000193
Then for any given bounded initial state x (0), for any given positive real number, a positive real number f can be foundminAnd TminSatisfies the following conditions: f. ofmin≥max{fmin1,fmin2,fmin3,fmin4When f is equal to fi≥fmin(i ═ 1,2), then state x is bounded, and
Figure BDA0001533404850000194
|yiless than or equal to (i-1, 2); further if x (0) is 0, then
Figure BDA0001533404850000195
|yi|≤(i＝1,2)。
In order to prove the tracking performance of the control method of the aircraft, the present embodiment performs numerical simulation verification on the tracking control performance of the flight controller, and specifically, the linear state feedback controller (i.e., the flight controller) designed in the present embodiment includes H∞A state feedback controller (i.e., equation (8)) and a robust compensator (i.e., equation (14)) were numerically simulated in a Matlab/Simulink environment to verify validity. The values of the nominal parameters are shown in table 1 below:
TABLE 1
Figure BDA0001533404850000196
Figure BDA0001533404850000201
H∞The parameters of the state feedback controller are given before, and the robust filtering parameters are selected as follows: f. of1＝50，f 25. High speed aircraft follows through speed and altitude passage in trim stateThe following two reference signals:
Figure BDA0001533404850000202
wherein r isvicAnd rhicIs a step reference input command, βrv＝0.6，βrh0.3. aircraft climbs from 110000 feet to 111000 feet in altitude, speed from 15060 feet/second to 15260 feet/second initial throttle opening command and angle of attack are β respectively0＝0.1762，α01.7905 degrees. In the simulation, all uncertainty parameters, aircraft parameters and aerodynamic coefficients were set to 50% of the nominal parameters. External disturbance di(i ═ v, γ, h, α, q, β) is given by the value dv＝6sin(0.2πt)+1,dγ＝0.002sin(0.1πt)-0.02,dh＝2sin(0.3πt)+0.4,dα＝0.05sin(0.2πt)-1,dq＝0.1sin(0.2πt)+0.05,dβ＝0.01sin(0.3πt)+0.3。
The numerical simulation results are shown in fig. 5 to 9; wherein, FIG. 5 is H∞A tracking response graph of a speed channel and a height channel of the controller to a reference signal; FIG. 6 is a graph of the tracking response of the velocity channel and the altitude channel of the linear state feedback controller to a reference signal; FIG. 7 is H∞Response graphs of track angle, attack angle and roll angular velocity of the controller and the linear state feedback controller; FIG. 8 is a graph of H in the numerical simulation process∞The controller and the linear state feedback controller are used for generating an input value change schematic diagram of an aircraft power system; FIG. 9 is H∞The controller and linear state feedback controller compare the tracking error of the velocity channel and the height channel.
As can be seen from fig. 5 to 9, the flight controller (i.e., linear state feedback controller) provided in the present embodiment is compared with the conventional H∞A controller that does not ignore the external disturbance and considers the external disturbance to include a time-invariant portion and a sinusoidally time-variant portion.
The control method of the aircraft provided by the embodiment of the invention is realized based on the flight controller, and has the following beneficial effects:
(1) the embodiment of the invention designs a robust linear controller aiming at a high-speed aircraft with strong nonlinearity and uncertainty, completely considers various uncertainties possibly involved in the current high-speed aircraft in the design process of the controller, and can well inhibit the influence of the uncertainties on the high-speed aircraft
(2) Example of the invention is in Standard H∞A robust compensator is designed on the basis of the controllers, the two controllers are combined to inhibit the influence of nonlinearity and uncertainty on a closed-loop system in the whole frequency domain, the reference signals of a speed channel and a height channel of the aircraft are tracked, and the tracking error converges in the prior neighborhood of an origin.
(3) The embodiment of the invention proves the proposed flight control method, and the method can be easily realized in practical application.
Corresponding to the above method embodiment, refer to the schematic structural diagram of a control device of an aircraft shown in fig. 10; the device is arranged on a controller of the aircraft; the device includes:
the signal acquisition module 10 is used for acquiring initial flight state parameters output by a power system of the aircraft; wherein the initial flight state parameters comprise flight speed and flight altitude;
the first signal generating module 11 is configured to generate an initial control signal according to the initial flight state parameter and a preset reference signal;
the error signal calculation module 12 is configured to input the initial control signal to a power system of the aircraft, acquire a current flight state parameter output by the power system, and calculate an error signal between the current flight state parameter and the reference signal;
a second signal generating module 13, configured to generate a compensation control signal according to the error signal;
and the control module 14 is used for determining throttle valve and rudder deflection angle control signals of the aircraft according to the initial control signal and the compensation control signal so as to control the flight state of the aircraft.
Further, the first signal generating module,and is also used for: calculating an initial control signal u by the following formulai,H：ui,H＝Ki,HeiI is 1, 2; wherein, i-1 represents the flying speed; i-2 represents the flying height; ki,HFeedback gain for suboptimal state; e.g. of the typeiIs an error signal.
Further, the first signal generating module is further configured to: suboptimal state feedback gain Ki,HObtained by the following formula:
Figure BDA0001533404850000221
wherein,
Figure BDA0001533404850000222
ωnis the natural angular frequency;
Figure BDA0001533404850000223
ρ is the atmospheric density; v. oftr i mThe speed of the aircraft in the balance state in the cruising flight stage; s is a reference area;
Figure BDA0001533404850000224
is the average aerodynamic chord length; cMeIs the aerodynamic coefficient; i isyyIs the moment of inertia; piIs a symmetric positive solution of the following equation:
Figure BDA0001533404850000225
Figure BDA0001533404850000226
Figure BDA0001533404850000227
a21＝vtrim；a22＝Ttrim/mN/Vtrim；CTβ0and CTβ2Is the aerodynamic coefficient; m is the aircraft mass; zetanIs the damping ratio; t istrimThrust of the aircraft in a balanced state in a cruising flight stage;
Figure BDA0001533404850000231
Figure BDA0001533404850000232
and thetaiIs a weight parameter; qi(i ═ 1,2) is a symmetric positive definite matrix, φi(i ═ 1,2) is the set attenuation factor; the superscript N indicates that the parameter is a nominal parameter.
According to the control device of the aircraft provided by the embodiment of the invention, an initial control signal can be generated according to the initial flight state parameter output by the aircraft and a preset reference signal; after the initial control signal is input into a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and a reference signal; then, according to the error signal, a compensation control signal can be generated; then according to the initial control signal and the compensation control signal, a throttle valve and a rudder deflection angle control signal of the aircraft can be determined, and the flight state of the aircraft is controlled; the method can inhibit the influence of various uncertain factors generated by the aircraft in the flight process by compensating the control signal, and improves the tracking control performance of the controller on the flight state of the aircraft.
Furthermore, the method adopts a linear time-invariant control method, is easy to realize and has strong practicability.
The embodiment of the invention also provides an aircraft, which comprises a processor and a sensor; the control device of the aircraft is arranged in the processor.
The aircraft provided by the embodiment of the invention has the same technical characteristics as the aircraft control method and device provided by the embodiment, so that the same technical problems can be solved, and the same technical effects can be achieved.
The aircraft control method and apparatus and the aircraft computer program product provided in the embodiments of the present invention include a computer-readable storage medium storing program codes, where instructions included in the program codes may be used to execute the methods described in the foregoing method embodiments, and specific implementations may refer to the method embodiments and are not described herein again.
The functions, if implemented in the form of software functional units and sold or used as a stand-alone product, may be stored in a computer readable storage medium. Based on such understanding, the technical solution of the present invention may be embodied in the form of a software product, which is stored in a storage medium and includes instructions for causing a computer device (which may be a personal computer, a server, or a network device) to execute all or part of the steps of the method according to the embodiments of the present invention. And the aforementioned storage medium includes: a U-disk, a removable hard disk, a Read-Only Memory (ROM), a Random Access Memory (RAM), a magnetic disk or an optical disk, and other various media capable of storing program codes.
Finally, it should be noted that: the above-mentioned embodiments are only specific embodiments of the present invention, which are used for illustrating the technical solutions of the present invention and not for limiting the same, and the protection scope of the present invention is not limited thereto, although the present invention is described in detail with reference to the foregoing embodiments, those skilled in the art should understand that: any person skilled in the art can modify or easily conceive the technical solutions described in the foregoing embodiments or equivalent substitutes for some technical features within the technical scope of the present disclosure; such modifications, changes or substitutions do not depart from the spirit and scope of the embodiments of the present invention, and they should be construed as being included therein. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.

Claims (3)

1. A method for controlling an aircraft, characterized in that it is applied to a controller of the aircraft; the method comprises the following steps:
acquiring initial flight state parameters output by a power system of an aircraft; wherein the initial flight state parameters include flight speed and flight altitude;
generating an initial control signal according to the initial flight state parameter and a preset reference signal;
inputting the initial control signal into a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and the reference signal;
generating a compensation control signal according to the error signal;
determining throttle valve and rudder deflection angle control signals of the aircraft according to the initial control signal and the compensation control signal so as to control the flight state of the aircraft;
the step of generating an initial control signal according to the initial flight state parameter and a preset reference signal comprises:
calculating an initial control signal u by the following formulai,H：
ui,H＝Ki,Hei,i＝1,2,；
Wherein, i-1 represents the flying speed; i-2 represents the flying height; ki,HFeedback gain for suboptimal state; e.g. of the typeiIs an error signal;
the suboptimal state feedback gain Ki,HObtained by the following formula:
Figure FDA0002491108560000011
wherein,
Figure FDA0002491108560000021
ωnis the natural angular frequency;
Figure FDA0002491108560000022
ρ is the atmospheric density; v. oftrimFor the aircraft in equilibrium during the cruising flight phaseSpeed; s is a reference area;
Figure FDA0002491108560000023
is the average aerodynamic chord length; cMeIs the aerodynamic coefficient; i isyyIs the moment of inertia;
Piis a symmetric positive solution of the following equation:
Figure FDA0002491108560000024
Figure FDA0002491108560000025
Figure FDA0002491108560000026
a21＝vtrim；a22＝Ttrim/mN/Vtrim；CTβ0and CTβ2Is the aerodynamic coefficient; m is the aircraft mass; zetanIs the damping ratio; t istrimThe thrust of the aircraft in the balance state in the cruising flight stage;
Figure FDA0002491108560000027
c1j，j＝1,2,3、c2jj is 1,2,3,4 and θiIs a weight parameter; qiI is 1,2 is a symmetric positive definite matrix, phiiI is 1,2 is a set attenuation factor; the superscript N represents that the parameter is a nominal parameter;
the step of generating a compensation control signal based on the error signal comprises:
calculating a compensation control signal u by the following formulai,R(s)：
Figure FDA0002491108560000028
Wherein s is a Laplace operator; f1(s)＝f1 3/(s+f1)3，
Figure FDA0002491108560000031
f1And f2Is a set filtering parameter which is a normal number; y is1＝v-rv，y2＝h-rh(ii) a v is the speed of the aircraft; h is the altitude of the aircraft; r isvThe reference speed of the aircraft in a cruising flight stage; r ishA reference altitude of the aircraft in a cruising flight phase; gi(s) is from ui,R(s) to yiThe transfer function of (a);
the transfer function Gi(s) obtained by the following formula:
Gi(s)＝Ci(sIi-Ai,H)-1Bi，i＝1,2；
wherein, IiI is 1,2 is an identity matrix;
Figure FDA0002491108560000032
Ai,H＝Ai+BiKi,H；
the step of determining throttle and rudder deflection angle control signals of the aircraft to control the flight state of the aircraft according to the initial control signal and the compensation control signal comprises:
calculating a final control signal u of the aircrafti＝ui,H+ui,R，i＝1,2；
And controlling the flying speed and/or flying height of the aircraft according to the final control signal.
2. A control device for an aircraft, characterized in that the device is provided to a controller for the aircraft; the device comprises:
the signal acquisition module is used for acquiring initial flight state parameters output by a power system of the aircraft; wherein the initial flight state parameters include flight speed and flight altitude;
the first signal generation module is used for generating an initial control signal according to the initial flight state parameter and a preset reference signal;
the error signal calculation module is used for inputting the initial control signal to a power system of the aircraft, acquiring current flight state parameters output by the power system, and calculating an error signal between the current flight state parameters and the reference signal;
the second signal generation module is used for generating a compensation control signal according to the error signal;
the control module is used for determining a throttle valve and a rudder deflection angle control signal of the aircraft according to the initial control signal and the compensation control signal so as to control the flight state of the aircraft;
the first signal generation module is further configured to:
calculating an initial control signal u by the following formulai,H：
ui,H＝Ki,Hei,i＝1,2,；
Wherein, i-1 represents the flying speed; i-2 represents the flying height; ki,HFeedback gain for suboptimal state; e.g. of the typeiIs an error signal;
the first signal generation module is further configured to: the suboptimal state feedback gain Ki,HObtained by the following formula:
Figure FDA0002491108560000041
wherein,
Figure FDA0002491108560000042
ωnis the natural angular frequency;
Figure FDA0002491108560000043
ρ is the atmospheric density; v. oftrimThe speed of the aircraft in the balance state in the cruising flight stage; s is a reference area;
Figure FDA0002491108560000044
is the average aerodynamic chord length; cMeIs the aerodynamic coefficient; i isyyIs the moment of inertia;
Piis a symmetric positive solution of the following equation:
Figure FDA0002491108560000045
Figure FDA0002491108560000046
Figure FDA0002491108560000051
a21＝vtrim；a22＝Ttrim/mN/Vtrim；CTβ0and CTβ2Is the aerodynamic coefficient; m is the aircraft mass; zetanIs the damping ratio; t istrimThe thrust of the aircraft in the balance state in the cruising flight stage;
Figure FDA0002491108560000052
c1j，j＝1,2,3、c2jj is 1,2,3,4 and θiIs a weight parameter; qiI is 1,2 is a symmetric positive definite matrix, phiiI is 1,2 is a set attenuation factor; the superscript N represents that the parameter is a nominal parameter;
the second signal generation module is further configured to:
calculating a compensation control signal u by the following formulai,R(s)：
Figure FDA0002491108560000053
Wherein s is a Laplace operator; f1(s)＝f1 3/(s+f1)3，
Figure FDA0002491108560000054
f1And f2Is a set filtering parameter which is a normal number; y is1＝v-rv，y2＝h-rh(ii) a v is the speed of the aircraft; h is the altitude of the aircraft; r isvThe reference speed of the aircraft in a cruising flight stage; r ishA reference altitude of the aircraft in a cruising flight phase; gi(s) is from ui,R(s) to yiThe transfer function of (a);
the transfer function Gi(s) obtained by the following formula:
Gi(s)＝Ci(sIi-Ai,H)-1Bi，i＝1,2；
wherein, IiI is 1,2 is an identity matrix;
Figure FDA0002491108560000055
Ai,H＝Ai+BiKi,H；
the control module is further configured to:
calculating a final control signal u of the aircrafti＝ui,H+ui,R，i＝1,2；
And controlling the flying speed and/or flying height of the aircraft according to the final control signal.
3. An aircraft, characterized in that the aircraft comprises a processor and a sensor; the apparatus of claim 2 disposed in the processor.
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