An aircraft overload control method based on angle of attack estimation

Abstract

The application provides an aircraft overload control method based on an angle of attack estimation, which firstly installs an accelerometer to measure a longitudinal overload signal and solves an overload-based angle of attack estimation signal according to aerodynamic parameter nominal values, then measures an angular rate signal through a pitch rate gyroscope, and designs a second-order differential filter network to obtain a pitch angle acceleration estimation signal, and then solves an angular rate-based angle of attack estimation signal by using the aerodynamic parameter nominal values; then a large time constant inertial integrator is designed according to an overload error signal to obtain an angle of attack and pitch angle rate related parameter adaptive estimation signal; and a virtual backstepping signal of a rudder deflection angle is composed of an equivalent control signal and an error feedback control signal; a second-order approximate differential signal of the rudder deflection angle is constructed, and saturation limiting, integration and initial value processing are performed to obtain a first-order differential signal of the rudder deflection angle, and finally a rudder deflection angle signal is inversely solved and transmitted to a pitch rudder system to realize tracking of an expected overload instruction signal of the aircraft.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of aircraft overload control, in particular to an aircraft overload control method based on angle of attack estimation. BACKGROUND

[0002] The angle of attack is the source of aircraft lift generation and the main factor of aircraft moment change; meanwhile, the rapid turning and high performance maneuver of the aircraft are closely related to the angle of attack, and the stability of the aircraft is also closely related to the angle of attack. Since the angle of attack exceeds a certain range, the aircraft will lose stability, and the range is small, and in the flight process, due to the uncertainty of high-speed air flow, it is difficult to accurately measure the angle of attack, so currently there is almost no application and implementation of measuring and introducing control and stability methods in any aircraft. The mainstream method at present is to measure the attitude angle, attitude angle rate and overload signal of the aircraft by using gyroscopes and accelerometers, and the accuracy of the measuring instrument is also high. Based on the above background reasons, we use the aerodynamic coefficients and overload signals and the attitude angle rate signals to reconstruct the angle of attack signal as the angle of attack estimation signal, and introduce it into the overload control to form the equivalent control term and the adaptive control term of the overload error feedback control; at the same time, the differential signal of the rudder deflection angle is compensated, and a high-quality overload tracking control method is obtained, which also makes the present application have high theoretical value and engineering value.

[0003] It should be noted that the information in the above background section of the invention is only used to strengthen the understanding of the background of the present application, and therefore can include information that does not constitute prior art known to those of ordinary skill in the art. SUMMARY

[0004] The purpose of the present application is to provide an aircraft overload control method based on angle of attack estimation, thereby overcoming the problem of low dynamic performance of aircraft overload control caused by the defects of related technologies.

[0005] According to one aspect of the present application, an aircraft overload control method based on angle of attack estimation is provided, comprising the following five steps:

[0006] Step S10, installing an accelerometer on the aircraft, measuring the longitudinal acceleration signal of the aircraft, and then converting it into a longitudinal overload signal; then introducing the elevator deflection angle signal of the aircraft and the flight speed signal of the aircraft and the nominal value of the aircraft air dynamic coefficient to obtain the overload estimation signal of the aircraft angle of attack; installing an angle measuring gyroscope on the aircraft, measuring the pitch angle rate signal of the aircraft, and then designing a second-order differential filter network to obtain the pitch angle acceleration approximation signal; then introducing the elevator deflection angle signal of the aircraft and the pitch angle rate signal and the nominal value of the aircraft air dynamic coefficient to obtain the angular acceleration estimation signal of the aircraft angle of attack through conversion; then performing average calculation on the overload estimation signal of the aircraft angle of attack and the angular acceleration estimation signal of the aircraft angle of attack to obtain the aircraft angle of attack estimation signal.

[0007] Step S20, according to the flight task of the pitch channel of the aircraft, setting the longitudinal overload expected signal, and comparing it with the longitudinal overload signal to obtain the longitudinal overload error signal; then performing integration to obtain the longitudinal overload error integral signal; then using a second-order differential filter network to obtain the longitudinal overload error approximate differential signal according to the longitudinal overload error signal; then designing an angle of attack related parameter adaptive estimation law according to the longitudinal overload error signal and the aircraft angle of attack estimation signal; then designing a large time constant inertial integrator to obtain the angle of attack related parameter adaptive estimation signal; then designing a pitch angle rate related parameter adaptive estimation law according to the longitudinal overload error signal and the pitch angle rate signal; then designing a large time constant inertial integrator to obtain the pitch angle rate related parameter adaptive estimation signal.

[0008] Step S30, constructing a deflection angle equivalent control signal according to the pitch angle acceleration approximation signal and the aircraft angle of attack estimation signal; then constructing a deflection angle error feedback control signal according to the longitudinal overload error signal, the longitudinal overload error integral signal and the longitudinal overload error approximate differential signal; then constructing a deflection angle adaptive estimation signal according to the angle of attack related parameter adaptive estimation signal and the pitch angle rate related parameter adaptive estimation signal; then superimposing the above-mentioned deflection angle equivalent control signal, the deflection angle error feedback control signal and the deflection angle adaptive estimation signal to obtain a deflection angle virtual backstepping signal.

[0009] Step S40, constructing a deflection angle second-order differential original signal according to the pitch angle acceleration approximation signal, the aircraft angle of attack estimation signal, the longitudinal overload error signal and the longitudinal overload error approximate differential signal; constructing a deflection angle first-order differential original signal according to the longitudinal overload error approximate differential signal; then designing a second-order differential filter and a first-order differential filter respectively, and then superimposing the deflection angle second-order differential original signal through the second-order differential filter and the deflection angle first-order differential original signal through the first-order differential filter to obtain a deflection angle second-order approximate differential signal.

[0010] Step S50, according to the rudder angle second-order approximation differential signal, saturation limiting is carried out to obtain a rudder angle second-order differential limiting signal; integration is carried out to obtain a rudder angle first-order differential signal; the initial value of the rudder angle first-order differential signal is set to 0 to obtain a rudder angle differential approximation compensation signal; and the rudder angle virtual inversion signal is superimposed to calculate a final pitch rudder angle signal, which is transmitted to a vehicle pitch channel rudder system to realize tracking of a desired overload command by a vehicle pitch channel longitudinal overload.

[0011] In an example embodiment of the present application, an accelerometer is installed on a vehicle to measure a longitudinal acceleration signal of the vehicle, which is then converted into a longitudinal overload signal; a pitch rudder angle signal of the vehicle and a flight speed signal of the vehicle and a nominal value of an air dynamic force coefficient of the vehicle are introduced to obtain an overload estimation signal of an attack angle of the vehicle; an angle rate gyroscope is installed on the vehicle to measure a pitch angle rate signal of the vehicle, and a second-order differential filter network is designed to obtain a pitch angle acceleration approximation signal; the pitch rudder angle signal of the vehicle and the pitch angle rate signal of the vehicle and the nominal value of the air dynamic force coefficient of the vehicle are introduced to obtain an angular acceleration estimation signal of the attack angle of the vehicle through conversion; and the overload estimation signal of the attack angle of the vehicle and the angular acceleration estimation signal of the attack angle of the vehicle are averaged to obtain an attack angle estimation signal of the vehicle, including:

[0012] ;

[0013] ;

[0014] ;

[0015] ;

[0016] ;

[0017] wherein is the longitudinal acceleration signal of the vehicle; is a gravity acceleration signal, which is a constant parameter; is the longitudinal overload signal; is the vehicle speed signal; is the pitch rudder angle signal, and the initial value thereof is set to 0; , is the nominal value of the air dynamic force coefficient of the vehicle, which is a constant parameter; is the overload estimation signal of the attack angle of the vehicle; is the pitch angle rate signal, , , , is a constant parameter of the second-order differential filter network; is a differential operator of a transfer function of a second-order differential filter network; is a pitch angle acceleration approximate signal; 、 、 is a nominal value of an air dynamic coefficient of the aircraft, and is a constant parameter; is an angle acceleration estimation signal of an attack angle of the aircraft; is an attack angle estimation signal of the aircraft.

[0018] In an example embodiment of the present application, according to a flight task of a pitch channel of the aircraft, a longitudinal overload expected signal is set, and compared with a longitudinal overload signal to obtain a longitudinal overload error signal; then integration is performed to obtain a longitudinal overload error integral signal; then a second-order differential filter network is used to obtain a longitudinal overload error approximate differential signal according to the longitudinal overload error signal; then an attack angle related parameter adaptive estimation law is designed according to the longitudinal overload error signal and the attack angle estimation signal of the aircraft; then a large time constant inertial integrator is designed to obtain an attack angle related parameter adaptive estimation signal; then a pitch angle rate related parameter adaptive estimation law is designed according to the longitudinal overload error signal and a pitch angle rate signal; then a large time constant inertial integrator is designed to obtain a pitch angle rate related parameter adaptive estimation signal including:

[0019] ;

[0020] ;

[0021] ;

[0022] ;

[0023] ;

[0024] ;

[0025] ;

[0026] wherein is the longitudinal overload expected signal, is the longitudinal overload error signal; is the longitudinal overload error integral signal; is a differential operator of a transfer function of a large time constant inertial integrator; is the longitudinal overload error approximate differential signal; is a constant parameter, is the attack angle related parameter adaptive estimation law; is a constant parameter of the large time constant inertial integrator, is the attack angle related parameter adaptive estimation signal; is a constant parameter for the pitch angle rate related parameter adaptive estimation law; is a constant parameter for the pitch angle rate related parameter adaptive estimation law; is a constant parameter for the large time constant inertial integrator, is a pitch angle rate related parameter adaptive estimation signal.

[0027] In an example embodiment of the present application, a rudder angle equivalent control signal is constructed according to the pitch angle acceleration approximation signal and the aircraft angle of attack estimation signal; a rudder angle error feedback control signal is constructed according to the longitudinal overload error signal, the longitudinal overload error integral signal, and the longitudinal overload error approximate differential signal; a rudder angle adaptive estimation signal is constructed according to the angle of attack related parameter adaptive estimation signal and the pitch angle rate related parameter adaptive estimation signal; and a rudder angle virtual inversion signal is obtained by superimposing the rudder angle equivalent control signal, the rudder angle error feedback control signal, and the rudder angle adaptive estimation signal, including:

[0028] ;

[0029] ;

[0030] ;

[0031] ;

[0032] wherein is a rudder angle equivalent control signal, , , is a constant control parameter for the error feedback control; is a rudder angle error feedback control signal, is a rudder angle adaptive estimation signal, is a rudder angle virtual inversion signal.

[0033] In an example embodiment of the present application, a rudder angle second-order differential original signal is constructed according to the pitch angle acceleration approximation signal, the aircraft angle of attack estimation signal, the longitudinal overload error signal, and the longitudinal overload error approximate differential signal; a rudder angle first-order differential original signal is constructed according to the longitudinal overload error approximate differential signal; a second-order differential filter and a first-order differential filter are designed respectively; the rudder angle second-order differential original signal is passed through the second-order differential filter; the rudder angle first-order differential original signal is passed through the first-order differential filter; and a rudder angle second-order approximate differential signal is obtained by collecting the signals after passing through the filters, including:

[0034] ;

[0035] ;

[0036] ;

[0037] wherein is a rudder angle second-order differential original signal; is a rudder angle first-order differential original signal; is a rudder angle second-order approximate differential signal; , is a constant time parameter of the second-order differential filter; is a constant time parameter of the first-order differential filter.

[0038] In an example embodiment of the present application, saturation limiting is performed according to the rudder angle second-order approximate differential signal to obtain a rudder angle second-order differential limited signal; integration is performed to obtain a rudder angle first-order differential signal; the initial value of the rudder angle first-order differential signal is set to 0 to obtain a rudder angle differential approximate compensation signal; and the rudder angle virtual inversion signal is superimposed to obtain a final pitch rudder angle signal including:

[0039] ;

[0040] ;

[0041] ;

[0042] ;

[0043] wherein is a rudder angle second-order differential limited signal; is a constant saturation parameter; is a constant integration parameter, is a rudder angle first-order differential signal; is a rudder angle differential approximate compensation signal, is a pitch rudder angle signal.

[0044] Advantageous effects

[0045] The application provides an aircraft overload control method based on an angle-of-attack estimation, and the main innovations are as follows: the first innovation is that, in view of the problem that the angle of attack of an aircraft is difficult to accurately measure, a method of using overload measurement and aerodynamic coefficient inverse solution is proposed to estimate the angle of attack of the aircraft; meanwhile, the pitch angle rate is measured by using a gyroscope, the angular acceleration approximate signal is obtained through differential filtering, and the aircraft angle of attack is estimated in combination with the aerodynamic coefficient, and finally the two average values are taken as the angle of attack estimation value. The method can reasonably utilize the measurement information of the accelerometer and the gyroscope, make a reasonable estimation of the angle of attack of the aircraft, and introduce the angle of attack as a key signal into the overload control to improve the dynamic characteristics of the overload control. The second innovation is that a large time constant inertial integrator is designed according to the overload error signal, the angle of attack and pitch angle rate related parameter adaptive estimation signal is obtained, and the rudder angle virtual backstepping signal is formed in combination with the rudder angle equivalent control signal and the rudder angle error feedback control signal, so that the error feedback and adaptive control can be reasonably matched, and the system uncertainty problem caused by the inaccurate aerodynamic parameters in the design can be compensated. The third innovation is that the rudder angle first-order differential signal is obtained through a second-order differential filter, a first-order differential filter, saturation limiting, integration and initial value processing, so that the rudder angle virtual backstepping signal is corrected, the non-minimum phase problem in the overload control is solved, and the dynamic quality of the overload control is further improved.

[0046] It should be understood that the foregoing general description and the following detailed description are only exemplary and explanatory, and cannot limit the application. BRIEF DESCRIPTION OF DRAWINGS

[0047] The drawings incorporated into the specification and forming a part of the specification, show embodiments consistent with the application, and together with the specification serve to explain the principles of the application. Obviously, the drawings in the following description are only some embodiments of the application, and other drawings can be obtained from these drawings without creative labor for those skilled in the art.

[0048] Figure 1 is a method flow chart of an aircraft overload control method based on an angle-of-attack estimation provided by the application;

[0049] Figure 2 is a longitudinal overload signal curve (unit: g) of the method provided by the embodiment of the application;

[0050] Figure 3 is a pitch angle rate signal curve (unit: degree per second) of the method provided by the embodiment of the application;

[0051] Figure 4 is an aircraft angle of attack estimation signal curve (unit: degree) of the method provided by the embodiment of the application;

[0052] Figure 5 is the longitudinal overload error signal curve (unit: g) of the method provided by the embodiment of the application;

[0053] Figure 6 is the longitudinal overload error approximate differential signal (unitless) of the method provided by the embodiment of the application;

[0054] Figure 7 is the rudder angle signal (unit: degree) of the method provided by the embodiment of the application. DETAILED DESCRIPTION

[0055] Example implementations will now be described more fully with reference to the accompanying drawings. Example implementations may, however, be implemented in many different forms and should not be construed as limited to the implementations set forth herein; rather, these implementations are provided so that this disclosure will be thorough and complete, and will fully convey the inventive aspects to those skilled in the art. The described features, structures, or characteristics can be combined in one or more implementations. In the following description, numerous specific details are provided to give a thorough understanding of implementations of the application. One skilled in the relevant art will recognize, however, that the aspects of the application can be practiced without one or more of the specific details, or with other methods, components, materials, and so forth. In other instances, well-known structures have not been described in detail so as not to obscure aspects of the application.

[0056] The application provides a flight vehicle overload control method based on angle of attack estimation, which first installs an accelerometer to measure a longitudinal overload signal and calculates an overload-based angle of attack estimation signal according to nominal values of aerodynamic parameters, then measures an angular rate signal through a pitch rate gyroscope, and designs a second-order differential filter network to obtain a pitch angle acceleration estimation signal, and then calculates an angular rate-based angle of attack estimation signal using the nominal values of the aerodynamic parameters; further, a large time constant inertial integrator is designed according to an overload error signal to obtain an angle of attack and pitch angle rate related parameter adaptive estimation signal; and a rudder angle virtual backstepping signal is composed of an equivalent control signal and an error feedback control signal; a rudder angle second-order approximate differential signal is constructed, and saturation limiting, integration and initial value processing are performed to obtain a rudder angle first-order differential signal, and finally the rudder angle signal is inversely solved and delivered to a pitch rudder system to realize tracking of a desired overload instruction signal by the flight vehicle.

[0057] Below, the application will be described in further detail with reference to the accompanying drawings. Figure 1 The application provides a flight vehicle overload control method based on angle of attack estimation, which first installs an accelerometer to measure a longitudinal overload signal and calculates an overload-based angle of attack estimation signal according to nominal values of aerodynamic parameters, then measures an angular rate signal through a pitch rate gyroscope, and designs a second-order differential filter network to obtain a pitch angle acceleration estimation signal, and then calculates an angular rate-based angle of attack estimation signal using the nominal values of the aerodynamic parameters; further, a large time constant inertial integrator is designed according to an overload error signal to obtain an angle of attack and pitch angle rate related parameter adaptive estimation signal; and a rudder angle virtual backstepping signal is composed of an equivalent control signal and an error feedback control signal; a rudder angle second-order approximate differential signal is constructed, and saturation limiting, integration and initial value processing are performed to obtain a rudder angle first-order differential signal, and finally the rudder angle signal is inversely solved and delivered to a pitch rudder system to realize tracking of a desired overload instruction signal by the flight vehicle. Figure 1 As shown in the figure, the flight vehicle overload control method based on angle of attack estimation comprises the following steps:

[0058] Step S10, specifically, can be divided into the following five small steps. First, install the accelerometer on the aircraft, measure the longitudinal acceleration signal of the aircraft, and then convert it to the longitudinal overload signal as follows:

[0059]

[0060] wherein is the longitudinal acceleration signal of the aircraft; is the gravity acceleration signal, which is a constant parameter, and in this case is selected as is the longitudinal overload signal, and in this case its variation curve is shown as Figure 2

[0061] Second, introduce the elevator deflection angle signal of the aircraft and the flight speed signal of the aircraft and the nominal value of the aircraft air dynamic force coefficient to obtain the overload estimation signal of the aircraft angle of attack as follows:

[0062]

[0063] wherein is the flight speed signal of the aircraft, which is a constant in this case, it is worth noting that the flight speed does not need to be accurately measured, only a rough estimate value is needed; is the elevator deflection angle signal, whose initial value is set to 0; is the nominal value of the aircraft air dynamic force coefficient, which is a constant parameter, and in this case is selected as is the overload estimation signal of the aircraft angle of attack.

[0064] Third, install the angle measuring gyroscope on the aircraft to measure the pitch angle rate signal of the aircraft, and then design a second-order differential filter network to obtain the pitch angle acceleration approximation signal as follows:

[0065]

[0066] wherein is the pitch angle rate signal, and its variation curve is shown as Figure 3 are constant parameters of the second-order differential filter network, and in this case are selected as is the differential operator of the transfer function of the second-order differential filter network; ​​​​​​​​​​​​​​​​The pitch angle acceleration approximation signal.

[0067] In the fourth step, the pitch rudder deflection angle signal and the pitch angle rate signal of the aircraft and the nominal value of the aircraft aerodynamic coefficients are introduced, and the angle acceleration estimation signal of the aircraft angle of attack is obtained by conversion as follows:

[0068] ;

[0069] Wherein , , is the nominal value of the aircraft aerodynamic coefficients, which is a constant parameter, and in this case it is selected as , , ; is the angle acceleration estimation signal of the aircraft angle of attack.

[0070] In the fifth step, the overload estimation signal of the aircraft angle of attack and the angle acceleration estimation signal of the aircraft angle of attack are averaged to obtain the aircraft angle of attack estimation signal as follows:

[0071] ;

[0072] Wherein is the aircraft angle of attack estimation signal, and in this case its transformation curve is shown in Figure 4 .

[0073] Step S20, which can be divided into the following five small steps. In the first step, according to the pitch channel flight task of the aircraft, the longitudinal overload expected signal is set, and compared with the longitudinal overload signal to obtain the longitudinal overload error signal as follows:

[0074] ;

[0075] Wherein is the longitudinal overload expected signal, and in this case it is selected as , is the longitudinal overload error signal, and its change curve is shown in Figure 5 .

[0076] In the second step, the longitudinal overload error signal is integrated to obtain the longitudinal overload error integral signal as follows:

[0077] ;

[0078] Wherein is the longitudinal overload error integral signal.

[0079] In the third step, the longitudinal overload error signal is further processed by a second-order differential filter network to obtain the longitudinal overload error approximate differential signal as follows:

[0080] ;

[0081] wherein is the longitudinal overload error approximate differential signal, the variation curve of which is shown in Figure 6 .

[0082] Fourthly, an angle of attack related parameter adaptive estimation law is designed according to the longitudinal overload error signal and the aircraft angle of attack estimation signal, and a large time constant inertial integrator is designed to obtain an angle of attack related parameter adaptive estimation signal as follows:

[0083] ;

[0084] ;

[0085] wherein is a constant parameter, which is selected as , is the angle of attack related parameter adaptive estimation law, which is selected as is a constant parameter of the large time constant inertial integrator, which is selected as is the angle of attack related parameter adaptive estimation signal.

[0086] Fifthly, a pitch rate related parameter adaptive estimation law is designed according to the longitudinal overload error signal and the pitch rate signal, and a large time constant inertial integrator is designed to obtain a pitch rate related parameter adaptive estimation signal as follows:

[0087] ;

[0088] ;

[0089] wherein is a constant parameter, which is selected as , is the pitch rate related parameter adaptive estimation law, is a constant parameter of the large time constant inertial integrator, which is selected as , is the pitch rate related parameter adaptive estimation signal.

[0090] Step S30 can be specifically divided into the following four sub-steps. Firstly, a rudder deflection angle equivalent control signal is constructed according to the pitch acceleration approximate signal and the aircraft angle of attack estimation signal as follows:

[0091] ;

[0092] wherein is the rudder deflection angle equivalent control signal.

[0093] Second, according to the longitudinal overload error signal, longitudinal overload error integral signal, longitudinal overload error approximate differential signal, the rudder angle error feedback control signal is constructed as follows:

[0094] ;

[0095] Wherein , , is a constant control parameter of error feedback control, and in the present case, it is selected as , , , is the rudder angle error feedback control signal.

[0096] Third, according to the angle of attack related parameter adaptive estimation signal, the pitch angle rate related parameter adaptive estimation signal, the rudder angle adaptive estimation signal is constructed as follows:

[0097] ;

[0098] Wherein is the rudder angle adaptive estimation signal.

[0099] Fourth, the above-mentioned rudder angle equivalent control signal, rudder angle error feedback control signal and rudder angle adaptive estimation signal are superimposed to obtain the rudder angle virtual inversion signal as follows:

[0100] ;

[0101] Wherein is the rudder angle virtual inversion signal.

[0102] Step S40, specifically, can be divided into the following three small steps. First, according to the pitch angle acceleration approximate signal, the aircraft angle of attack estimation signal, the longitudinal overload error signal, the longitudinal overload error approximate differential signal, the rudder angle second-order differential original signal is constructed as follows:

[0103] ;

[0104] Wherein is the rudder angle second-order differential original signal.

[0105] Second, according to the longitudinal overload error approximate differential signal, the rudder angle first-order differential original signal is constructed as follows:

[0106] ;

[0107] Wherein is the rudder angle first-order differential original signal.

[0108] Third step, respectively, design a second-order differential filter and a first-order differential filter, the rudder angle second-order differential original signal through the second-order differential filter, the rudder angle first-order differential original signal through the first-order differential filter after the collection of the rudder angle second-order approximate differential signal as follows:

[0109]

[0110] Wherein is the rudder angle second-order approximate differential signal; , is the constant time parameter of the second-order differential filter; is the constant time parameter of the first-order differential filter, in this case, selected as , , .

[0111] Step S50, specifically, can be divided into the following four small steps. First, according to the rudder angle second-order approximate differential signal saturation limiting, the rudder angle second-order differential limiting signal as follows:

[0112]

[0113] Wherein is the rudder angle second-order differential limiting signal; is the constant saturation parameter; is the constant integral parameter, in this case, selected as , .

[0114] Second, according to the rudder angle second-order differential limiting signal integral, the rudder angle first-order differential signal as follows:

[0115]

[0116] Wherein is the rudder angle first-order differential signal.

[0117] Third, set the initial value to 0, the rudder angle differential approximate compensation signal as follows:

[0118]

[0119] Wherein is the rudder angle differential approximate compensation signal.

[0120] Fourth, superimposed on the rudder angle virtual inversion signal, the transformation processing is calculated out of the final pitch rudder angle signal as follows:

[0121] ;​​​​

[0122] in The pitch deflection signal, in this case, its variation curve is as follows: Figure 7 As shown.

[0123] Finally, the pitch deflection signal is transmitted to the aircraft's pitch channel rudder system, enabling the aircraft's pitch channel longitudinal overload to track the desired overload command.

[0124] Depend on Figure 2 It can be seen that the final aircraft overload can stably track the desired command 2 in about 0.8 seconds, indicating that the method provided by this invention has excellent speed. Figure 5 It can be seen that the overload error can quickly converge to 0, and the steady-state error is very small, which shows that the method provided by the present invention has good accuracy. Figure 3 The pitch rate curve is shown. Figure 4 The angle of attack estimation curve shows its relationship with the overload variation curve, i.e. Figure 2 It has good matching and synchronization characteristics and conforms to the physical characteristics of the aircraft. Since the overload is mainly caused by the angle of attack, it can be seen that the angle of attack estimation has good dynamic characteristics. Figure 7 The final pitch deflection curve shows a smooth change without spikes or peaks, and the maximum value does not exceed 4 degrees, meeting the requirements for engineering implementation. In summary, the method provided by this invention exhibits excellent overload stability tracking dynamic performance and good dynamic performance in angle of attack estimation, thus making the method of this invention highly valuable in both theory and engineering practice.

Claims

1. An aircraft overload control method based on angle of attack estimation, characterized in that, The method comprises the following steps: Step S10, installing an accelerometer on the aircraft, measuring the longitudinal acceleration signal of the aircraft, and then converting it into a longitudinal overload signal; then introducing the elevator deflection angle signal of the aircraft and the flight speed signal of the aircraft and the nominal value of the aircraft air dynamic force coefficient to obtain an angle of attack overload estimation signal of the aircraft; installing an angle measuring gyroscope on the aircraft, measuring the pitch angle rate signal of the aircraft, and then designing a second-order differential filter network to obtain an approximate pitch angle acceleration signal; then introducing the elevator deflection angle signal of the aircraft and the pitch angle rate signal and the nominal value of the aircraft air dynamic force coefficient to obtain an angular acceleration estimation signal of the angle of attack of the aircraft through conversion; then performing average calculation on the overload estimation signal of the angle of attack of the aircraft and the angular acceleration estimation signal of the angle of attack of the aircraft to obtain an angle of attack estimation signal of the aircraft as follows: ； ； ； ； ； wherein is a longitudinal acceleration signal of the aircraft; is a gravitational acceleration signal, taken as a constant parameter; is a longitudinal overload signal; is a speed signal of the aircraft; is a pitch control deflection signal, whose initial value is set to 0; , is a nominal value of the aerodynamic coefficients of the aircraft, taken as a constant parameter; is an overload estimation signal of the angle of attack of the aircraft; is a pitch rate signal, , , , is a constant parameter of the second-order derivative filter network; is a differential operator of the transfer function of the second-order derivative filter network; is a pitch acceleration approximation signal; , , is a nominal value of the aerodynamic coefficients of the aircraft, taken as a constant parameter; is an angular acceleration estimation signal of the angle of attack of the aircraft; is an angle of attack estimation signal of the aircraft; Step S20, setting a longitudinal overload expected signal according to the flight task of the pitch channel of the aircraft, and comparing it with the longitudinal overload signal to obtain a longitudinal overload error signal; Then, the integral is obtained, and the longitudinal overload error integral signal is obtained; then, the second-order differential filter network is adopted to obtain the approximate differential signal of the longitudinal overload error according to the longitudinal overload error signal; then, the pitch angle related parameter adaptive estimation law is designed according to the longitudinal overload error signal and the angle of attack estimation signal of the aircraft; then, the large time constant inertial integrator is designed to obtain the pitch angle related parameter adaptive estimation signal as follows: ； ； ； ； ； ； ； wherein is a longitudinal overload desired signal, is a longitudinal overload error signal; is a longitudinal overload error integral signal; is a differential operator of a large time constant inertial integrator transfer function; is a longitudinal overload error approximate differential signal; is a constant parameter, is an angle of attack dependent parameter adaptive estimation law; is a constant parameter of a large time constant inertial integrator, is an angle of attack dependent parameter adaptive estimation signal; is a constant parameter, is a pitch rate dependent parameter adaptive estimation law; is a constant parameter of a large time constant inertial integrator, is a pitch rate dependent parameter adaptive estimation signal; Step S30, constructing an equivalent control signal of the deflection angle according to the approximate pitch angle acceleration signal and the angle of attack estimation signal of the aircraft; then, the deflection angle error feedback control signal is constructed according to the longitudinal overload error signal, the longitudinal overload error integral signal and the approximate differential signal of the longitudinal overload error; then, the deflection angle adaptive estimation signal is constructed according to the pitch angle related parameter adaptive estimation signal and the pitch angle rate related parameter adaptive estimation signal; then, the deflection angle virtual backstepping signal is obtained by superimposing the above-mentioned equivalent control signal of the deflection angle, the deflection angle error feedback control signal and the deflection angle adaptive estimation signal as follows: ； ； ； ； wherein is a rudder deflection angle equivalent control signal, , , is a constant control parameter of error feedback control; is a rudder deflection angle error feedback control signal, is a rudder deflection angle adaptive estimation signal, is a rudder deflection angle virtual inversion signal; Step S40, constructing a second-order differential original signal of the deflection angle according to the approximate pitch angle acceleration signal, the angle of attack estimation signal of the aircraft, the longitudinal overload error signal and the approximate differential signal of the longitudinal overload error; constructing a first-order differential original signal of the deflection angle according to the approximate differential signal of the longitudinal overload error; Then, the second-order differential filter and the first-order differential filter are designed respectively, the second-order differential filter is used to process the second-order differential original signal of the deflection angle, the first-order differential filter is used to process the first-order differential original signal of the deflection angle, and the second-order approximate differential signal of the deflection angle is obtained after the signals are collected as follows: ； ； ； wherein is a second order differential of the rudder angle; is a first order differential of the rudder angle; is a second order approximate differential of the rudder angle; , is a constant time parameter of the second order differential filter; is a constant time parameter of the first order differential filter; Step S50, according to the rudder angle second order approximation differential signal, saturation limiting is carried out to obtain the rudder angle second order differential limiting signal; then integration is carried out to obtain the rudder angle first order differential signal; then the initial value is set to 0 to obtain the rudder angle differential approximation compensation signal; then the rudder angle virtual inversion signal is superimposed to calculate the final pitch rudder angle signal, which is transmitted to the aircraft pitch channel rudder system to realize the tracking of the expected overload command by the aircraft pitch channel longitudinal overload as follows: ； ； ； ； wherein is a rudder angle second order derivative limiter signal; is a constant saturation parameter; is a constant integration parameter, is a rudder angle first order derivative signal; is a rudder angle derivative approximate compensation signal, is a pitch rudder angle signal.

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