Aircraft robust control method based on angular acceleration feedback

Abstract

This invention provides a robust control method for aircraft based on angular acceleration feedback, comprising the following steps: First, constructing the three-axis attitude angle control law for the aircraft and writing the nonlinear dynamic equations of the aircraft; then, designing control signals considering the flight quality of the aircraft using the inverse model method, and using angular acceleration signals to replace model parameters to reduce the dependence of the control method on the model. However, considering the uncertainty of the aircraft model, the parameters considering flight quality need to be optimized. Taking pitch control as an example, a longitudinal closed-loop control structure considering angular acceleration feedback is proposed. A longitudinal controller is designed based on a proportional-integral controller. In control, only the gain of the command path pre-filter needs to be adjusted through the angular acceleration error signal to ensure the effectiveness of the angular acceleration feedback, thus achieving good longitudinal tracking performance and solving the problem of poor applicability of traditional torque commands for aircraft control.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aircraft control method, and particularly relates to an aircraft robust control method based on angular acceleration feedback. BACKGROUND

[0002] The acceleration signal can directly reflect the influence of external force disturbance on the object, and good control of the angular acceleration signal can adjust the object in the disturbed state back to the stable state. For a system containing nonlinear and coupling problems, the two problems usually act on the angular acceleration of the system first. Therefore, good control of the angular acceleration signal is of great significance to the control problem of such a system. The introduction of angular acceleration feedback in the servo system can make the system have certain robustness to the change of torque and load moment of inertia, certain robustness to external disturbance and load change, and can also suppress resonance and nonlinear friction. Therefore, the introduction of angular acceleration feedback is an effective method to improve the performance of the system.

[0003] At present, acceleration feedback is mostly applied to robot control or vibration control of mechanical hands. The current research results show that it can improve the dynamic stiffness of the system, increase the bandwidth of the system, and enhance the stability of the system. In particular, the angular acceleration signal feedback in the inverse model control method of the aircraft can eliminate the feedback term containing the state, and the angular acceleration signal replaces the independent aerodynamic model parameters, so that the control law is not affected by the model information such as static derivative and damping, thereby reducing the dependence of the control method on the model and improving the robustness of the system. Moreover, compared with the torque command, the angular acceleration control command can unify the commands of aircrafts of different sizes and masses, thereby improving the applicability of the control technology.

[0004] Based on this, an aircraft robust control method based on angular acceleration feedback is proposed. SUMMARY

[0005] The technical problem to be solved by the application is to provide an aircraft robust control method based on angular acceleration feedback to solve the problems in the background art.

[0006] To solve the above technical problems, the technical solution adopted by the application is that an aircraft robust control method based on angular acceleration feedback comprises the following steps:

[0007] S1, constructing a three-axis attitude angle control law of the aircraft;

[0008] S2, defining a nonlinear dynamics equation of the aircraft, and based on the small disturbance assumption, separating the nonlinear and linear parts of the equation and representing them as follows:

[0009] ;

[0010] where x is the state variable, is the derivative of the state variable, u is the input variable, f is a nonlinear state variable dynamics function, and l is a linear function;

[0011] The control signal containing the desired flight quality is defined using the inverse model method, the control signal is expressed as two parts of the previous control signal and the incremental control command, and the control signal is brought into the nonlinear dynamics equation of the aircraft to eliminate the dynamics characteristics of the aircraft;

[0012] S3, taking pitch control as an example, the longitudinal dynamics structure is divided into three parts of inner loop control, feedforward control loop and command path pre-filter, the longitudinal dynamics is based on proportional integral controller, and the angular acceleration and pitch angular velocity are used as feedback variables, and the feedforward control loop and the command path pre-filter are used to improve the initial pitch angular acceleration and the control quality during maneuvering.

[0013] Further, in S1, the three-axis attitude angle control law of the aircraft adopts a three-loop cascade control structure, the outermost loop is an attitude angle control loop, and the attitude angle tracking error is used to generate a desired attitude angle rate command through an attitude angle controller;

[0014] The middle loop is an attitude angle rate control, which uses the attitude angle rate command generated by the outermost loop to generate an attitude angle acceleration command through an attitude angle rate controller, and the command is transmitted to the innermost loop angular acceleration controller;

[0015] The innermost loop is an attitude angle acceleration control, which uses the estimated angular acceleration signal for feedback, and simultaneously acts with the attitude angle acceleration command generated by the middle loop controller to generate the aircraft control input signal.

[0016] Further, in S2, the control command is defined as the sum of the control command at the previous moment and the incremental control command , and is written in the following form:

[0017] ;

[0018] ;

[0019] Let be reversible, then the control law can be obtained as follows:

[0020] ;

[0021] In the formula, is the model-based angular acceleration estimation;

[0022] Accordingly, Let the desired state of the flight quality be the rate of reaching the design, and the variables of the above equation are replaced to obtain the following equation:

[0023] ;

[0024] In the formula, is the angular acceleration expected signal containing the desired flight quality;

[0025] is the angular acceleration signal obtained based on the model, that is, the actual value of the aircraft angular acceleration signal;

[0026] The previous control signal and the incremental control command are combined to design the current control instruction, and the control signal at the current time It can be designed as:

[0027] ;

[0028] And the equation of the signal will be The equation of the aircraft nonlinear dynamics equation is obtained as follows:

[0029] .

[0030] Further, the longitudinal motion equation of the aircraft can be written as follows:

[0031] ;

[0032] Where, p is the roll angular velocity of the aircraft;

[0033] q is the pitch angular velocity of the aircraft;

[0034] r is the yaw angular velocity of the aircraft, with the unit of ,

[0035] is the pitch angular acceleration of the aircraft, with the unit of ,

[0036] I xx ,I yy ,I zz is the moment of inertia of the aircraft,

[0037] I xz is the product of the moment of inertia of the aircraft;

[0038] Assuming the longitudinal moment For the aerodynamic derivatives to be linear, written as follows:

[0039] ;

[0040] where, is the linearized angle of attack pitch moment;

[0041] is the linearized angular velocity pitch moment;

[0042] is the linearized elevator deflection pitch moment;

[0043] is the rudder deflection angle, is the angle of attack of the aircraft;

[0044] Combining the above two equations, the equation combining both linear and nonlinear parts can be obtained as:

[0045] ;

[0046] In the above equation, represents the linearized moment, defined as:

[0047] ;

[0048] Finally, reversing the above equation gives the control law based on the inverse model method:

[0049] ;

[0050] where, is the calculated pitch angle acceleration according to the desired dynamics, is the calculated rudder deflection angle, is the measured value of the angle of attack on the aircraft, is the measured value of the roll angular velocity on the aircraft, and is the measured value of the pitch angular velocity on the aircraft, is the measured value of the yaw angular velocity on the aircraft.

[0051] Further, the desired longitudinal dynamics design is based on a proportional-integral controller with the structure of angle acceleration and pitch angular velocity as feedback variables, and the desired angle acceleration is represented as:

[0052] ;

[0053] where, is the command acceleration, is the normal acceleration, the proportional control parameter for the pitch angle acceleration, the integral control parameter for the pitch angle acceleration, the control parameter for the normal acceleration, the control parameter for the pitch angle velocity;

[0054] The short period mode of the longitudinal motion is expressed as

[0055] ;

[0056] where, the true velocity of the aircraft in units, the pitch angle time constant, the derivative of the angle of attack of the aircraft, the acceleration of gravity, the short period mode of the aircraft can be expressed as a transfer function the complex variable in the transfer function;

[0057] Combining the above two equations, the transfer function is obtained as follows:

[0058] ;

[0059] The characteristic equation of the closed loop system is:

[0060] In order to eliminate , is:

[0061] ;

[0062] In the above equation, is the damping ratio of the short period mode, is the natural frequency of the short period mode;

[0063] Then, the initial values of the flight quality parameters are obtained as follows:

[0064] ;

[0065] Also obtained:

[0066] ;

[0067] For the pitch angle velocity of the control command, it is written from the aircraft longitudinal equation as:

[0068] ;

[0069] Combining the above three equations, we have:

[0070] ;

[0071] Pole-zero cancellation is performed on the pitch angular velocity expression:

[0072] ;

[0073] In the formula, The desired pitch angle time constant selected according to flight quality;

[0074] An expression containing s;

[0075] for The pole-related expression obtained after transformation.

[0076] Furthermore, Related to command path pre-filter, The control parameters derived from the actual overload are used in the command path pre-filter. For the control parameters derived from command overload in the command path pre-filter:

[0077] .

[0078] Compared with the prior art, the present invention has the following advantages:

[0079] This invention first constructs a three-axis attitude angle control law for an aircraft, writes the nonlinear dynamic equations of the aircraft, and uses the inverse model method to design control signals that consider the flight quality of the aircraft. It uses angular acceleration signals to replace model parameters, reducing the dependence of the control method on the model. Considering the uncertainty of the aircraft model, the parameters considering flight quality need to be optimized. Taking pitch control as an example, a longitudinal closed-loop control structure considering angular acceleration feedback is proposed. The longitudinal controller is designed based on a proportional-integral controller. In the control, only the gain of the command path pre-filter needs to be adjusted through the angular acceleration error signal to ensure the effectiveness of angular acceleration feedback and achieve good longitudinal tracking performance. Attached Figure Description

[0080] Figure 1 This is a schematic diagram of the attitude angle control law architecture of the present invention;

[0081] Figure 2 This is a flowchart of the angular acceleration feedback control law based on dynamic inverse in an embodiment of the present invention;

[0082] Figure 3 This is a schematic diagram of the longitudinal closed-loop control structure in an embodiment of the present invention;

[0083] Figure 4 This is a comparison diagram of the control method in the embodiments of the present invention and the normal longitudinal control result without the introduction of angular acceleration feedback;

[0084] Figure 5 This is a comparison chart of the longitudinal control results of two control strategies under pulse interference in this embodiment of the invention;

[0085] Figure 6 The diagram shows the amplitude-frequency and phase-frequency characteristics of the control loop without angular acceleration feedback in this embodiment of the invention; the phase margin is 67.1° and the amplitude margin is 10.7dB.

[0086] Figure 7 The diagram shows the amplitude-frequency and phase-frequency characteristics of the control loop with angular acceleration feedback introduced in this embodiment of the invention. The phase margin is 77.1° and the amplitude margin is 16.2dB. Detailed Implementation

[0087] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0088] like Figures 1-3 As shown, the present invention provides a technical solution: a robust control method for aircraft based on angular acceleration feedback, comprising the following steps:

[0089] S1, the three-axis attitude angle control law of the architecture aircraft; specifically as follows: Figure 1 As shown,

[0090] The three-axis attitude angle control law of the aircraft is composed of a three-loop cascaded control. The outermost loop is the attitude angle control loop, which uses the attitude angle tracking error to generate the desired attitude angle rate command through the attitude angle controller.

[0091] The middle loop is for attitude angular rate control. It uses the attitude angular rate command generated by the outermost loop to generate the attitude angular acceleration command through the attitude angular rate controller and then transmits it to the angular acceleration controller of the innermost loop.

[0092] The innermost loop is for attitude angular acceleration control, which uses the estimated angular acceleration signal for feedback. This feedback, along with the attitude angular acceleration command generated by the intermediate loop controller, simultaneously produces the aircraft control input signal.

[0093] S2. Define the nonlinear dynamic equations of the aircraft, and based on the small perturbation assumption, express its nonlinear and linear parts separately:

[0094] ;

[0095] in, x It is a state variable. It is the derivative of the state quantity. u It is the input quantity. f For nonlinear state dynamic functions, l It is a linear function;

[0096] The inverse model method is used to define the control signal containing the desired flight quality. The control signal is expressed as two parts: the previous control signal and the incremental control command. Then, the control signal is substituted into the nonlinear dynamic equation of the aircraft to eliminate the dynamic characteristics of the aircraft.

[0097] Specifically, such as Figure 2 As shown, assuming control commands Defined as the control command of the previous moment With incremental control commands The sum of these can be written in the following form:

[0098] ;

[0099] ;

[0100] set up If it is reversible, then the following control law can be obtained:

[0101] ;

[0102] In the formula, This refers to model-based angular acceleration estimation;

[0103] Based on this Let the rate be the desired state for achieving the designed flight quality, and by substituting the variables in the above equation, we obtain the following equation:

[0104] ;

[0105] In the formula, The desired angular acceleration signal includes the desired flight characteristics;

[0106] This refers to the angular acceleration signal obtained based on the model, i.e., the actual value of the aircraft's angular acceleration signal;

[0107] By combining previous control signals and incremental control commands to design the current control instruction, the control signal at the current moment is... It can be designed as follows:

[0108] ;

[0109] And, will produce Substituting the signal equation into the nonlinear dynamics equation of the aircraft yields the following equation:

[0110] .

[0111] S3. Taking pitch control as an example, the longitudinal dynamics structure is divided into three parts: inner loop control, feedforward control loop and command path pre-filter. The longitudinal dynamics is based on a proportional-integral controller, with angular acceleration and pitch angular velocity as feedback variables. The feedforward control loop and command path pre-filter are used to improve the initial pitch angular acceleration and handling quality during maneuvers.

[0112] The longitudinal motion equation of the aircraft can be written in the following form:

[0113] ;

[0114] in, p The roll angular velocity of the aircraft;

[0115] q The pitch rate of the aircraft;

[0116] r The yaw rate of the aircraft, in units of . ,

[0117] The pitch acceleration of the aircraft, in units of . ,

[0118] I xx ,I yy ,I zz It is the moment of inertia of the aircraft.

[0119] I xz It is the product of the aircraft's moments of inertia;

[0120] Assuming longitudinal moment For the aerodynamic derivative to be linear, it can be written in the following form:

[0121] ;

[0122] in, It is the linearized pitch moment at the angle of attack;

[0123] It is the linearized angular velocity pitch torque;

[0124] It is the linearized elevator pitch moment;

[0125] The deflection angle of the horizontal tail. Angle of attack of the aircraft;

[0126] Combining the above two equations, we can derive an equation that combines the linear and nonlinear components:

[0127] ;

[0128] In the above formula The linearized torque is defined as follows:

[0129] ;

[0130] Finally, reversing the above equations yields the control law based on the inverse model method:

[0131] ;

[0132] in The pitch acceleration is calculated based on the desired dynamics. The calculated tail yaw angle, This refers to the measured angle of attack on the aircraft. This refers to the measured value of the roll angular velocity on the aircraft. and This refers to the measured pitch angular velocity of the aircraft. This is a measured value of the yaw rate on the aircraft.

[0133] Specifically, such as Figure 3 As shown, the desired longitudinal dynamics design is based on a proportional-integral controller with angular acceleration and pitch angular velocity as feedback variables. The desired angular acceleration is expressed as follows:

[0134] ;

[0135] in, For command acceleration, Normal acceleration, The proportional control parameter for pitch angle acceleration. The integral control parameter for pitch angle acceleration, These are the control parameters for normal acceleration. These are the control parameters for pitch angular velocity;

[0136] The short-period modes of longitudinal motion are represented as follows:

[0137] ;

[0138] in, For The actual speed of the aircraft is expressed in units of... The pitch angle time constant, Let be the derivative of the aircraft's angle of attack. For gravitational acceleration, the short-period modes of an aircraft can be represented by a transfer function. For complex variables in the transfer function;

[0139] Combining the above two equations, we can obtain the transfer function as follows;

[0140] ;

[0141] The characteristic equation of the closed-loop system is:

[0142] In order to eliminate , for:

[0143] ;

[0144] In the above formula, It is the damping ratio of the short-period mode. It is the natural frequency of the short-period mode;

[0145] Then, the initial values ​​of the flight quality parameters are obtained as follows:

[0146] ;

[0147] You can also get:

[0148] ;

[0149] The pitch rate for control commands can be written from the aircraft's longitudinal equations as:

[0150] ;

[0151] Combining the above three formulas, we have:

[0152] ;

[0153] Pole-zero cancellation is performed on the pitch angular velocity expression:

[0154] ;

[0155] In the formula, The desired pitch angle time constant selected according to flight quality;

[0156] An expression containing s;

[0157] for The pole-related expression obtained after transformation.

[0158] Related to command path pre-filter, The control parameters derived from the actual overload are used in the command path pre-filter. These are the control parameters derived from command overload in the command path pre-filter;

[0159] .

[0160] The above formula is a command path pre-filter. The gain of this filter can provide good overall acquisition performance within the flight envelope, but the overshoot of pitch rate will affect the longitudinal fine tracking performance.

[0161] To improve longitudinal fine-grained tracking performance without affecting overall acquisition performance, a gain scheduler based on the acceleration error signal was designed. The gain is the lever force. and actual lever force Functions between.

[0162] The proposed control method, under the same control architecture and acceleration response type, can provide the good response required for precise tracking tasks.

[0163] After the above steps, a robust control technology for aircraft based on angular acceleration feedback is provided. This control method is applied to the nonlinear dynamic equations of the aircraft established in S2. That is, by giving a pitch angle command signal, the tracking effect of the pitch angle is observed and compared with the normal longitudinal control without angular acceleration feedback to observe the control effect after introducing angular acceleration feedback.

[0164] Observe the pitch angle tracking effects of the two control strategies. Figure 4 This is a comparison diagram of the control method of the present invention and the normal longitudinal control results without the introduction of angular acceleration feedback; Figure 5 A comparison of the longitudinal control results of two control strategies under the introduction of impulse interference; Figure 6 The amplitude-frequency and phase-frequency characteristics of the control loop without angular acceleration feedback are shown; the phase margin is 67.1° and the amplitude margin is 10.7dB. Figure 7 The amplitude-frequency and phase-frequency characteristics of the control loop for introducing angular acceleration feedback are shown in the diagram. The phase margin is 77.1° and the amplitude margin is 16.2dB.

[0165] It can be seen that the control loop with angular acceleration feedback exhibits better response performance to step signals and pulse interference than the control loop without angular acceleration feedback. The control loop with angular acceleration feedback also has greater phase and amplitude margins than the control loop without angular acceleration feedback, effectively improving the system's technical capabilities and avoiding the problem of excessively high phase margins leading to implementation difficulties.

[0166] It should be noted that, in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article, or apparatus.

[0167] Although embodiments of the invention have been shown and described, it will be understood by those skilled in the art that various changes, modifications, substitutions and alterations can be made to these embodiments without departing from the principles and spirit of the invention, the scope of which is defined by the appended claims and their equivalents.

Claims

1. A robust control method for aircraft based on angular acceleration feedback, characterized in that: Includes the following steps: S1, the three-axis attitude angle control law of the architecture aircraft; S2. Define the nonlinear dynamic equations of the aircraft, and based on the small perturbation assumption, express its nonlinear and linear parts separately: ； Where x is a state variable, It is the derivative of the state quantity. u It is the input quantity. f For nonlinear state dynamic functions, l It is a linear function; The inverse model method is used to define the control signal containing the desired flight quality. The control signal is expressed as two parts: the previous control signal and the incremental control command. Then, the control signal is substituted into the nonlinear dynamic equation of the aircraft to eliminate the dynamic characteristics of the aircraft. S3. Taking pitch control as an example, the longitudinal dynamics structure is divided into three parts: inner loop control, feedforward control loop and command path pre-filter. The longitudinal dynamics is based on a proportional-integral controller, with angular acceleration and pitch angular velocity as feedback variables. The feedforward control loop and command path pre-filter are used to improve the initial pitch angular acceleration and handling quality during maneuvers. In S1, the three-axis attitude angle control law of the aircraft is composed of a three-loop cascaded control. The outermost loop is the attitude angle control loop, which uses the attitude angle tracking error to generate the desired attitude angle rate command through the attitude angle controller. The middle loop is for attitude angular rate control. It uses the attitude angular rate command generated by the outermost loop to generate the attitude angular acceleration command through the attitude angular rate controller and then transmits it to the angular acceleration controller of the innermost loop. The innermost loop is for attitude angular acceleration control, which uses the estimated angular acceleration signal for feedback. This feedback, along with the attitude angular acceleration command generated by the intermediate loop controller, simultaneously produces the aircraft control input signal.

2. The robust control method for aircraft based on angular acceleration feedback according to claim 1, characterized in that, In S2, assume the control command Defined as the control command of the previous moment With incremental control commands The sum of these can be written in the following form: ； ； set up If it is reversible, then the following control law can be obtained: ； In the formula, This refers to model-based angular acceleration estimation; Based on this Let the rate be the desired state for achieving the designed flight quality, and by substituting the variables in the above equation, we obtain the following equation: ； In the formula, The desired angular acceleration signal includes the desired flight characteristics; This refers to the angular acceleration signal obtained based on the model, i.e., the actual value of the aircraft's angular acceleration signal; By combining previous control signals and incremental control commands to design the current control instruction, the control signal at the current moment is... Designed as follows: ； And, will produce Substituting the signal equation into the nonlinear dynamics equation of the aircraft yields the following equation: 。 3. The robust control method for aircraft based on angular acceleration feedback according to claim 2, characterized in that, The longitudinal motion equation of the aircraft can be written in the following form: ； in, p The roll angular velocity of the aircraft; q The pitch rate of the aircraft; r The yaw rate of the aircraft, in units of . ; The pitch acceleration of the aircraft, in units of . ; I xx ,I yy ,I zz It is the moment of inertia of the aircraft. I xz It is the product of the aircraft's moments of inertia; Assuming longitudinal moment For the aerodynamic derivative to be linear, it can be written in the following form: ； in, It is the linearized pitch moment at the angle of attack; It is the linearized angular velocity pitching moment; It is the linearized elevator pitch moment; The deflection angle of the horizontal tail. Angle of attack of the aircraft; Combining the above two equations, we can derive an equation that combines the linear and nonlinear components: ； In the above formula The linearized torque is defined as follows: ； Finally, reversing the above equations yields the control law based on the inverse model method: ； in, The pitch acceleration is calculated based on the desired dynamics. The calculated tail yaw angle, This refers to the measured angle of attack on the aircraft. This refers to the measured value of the roll angular velocity on the aircraft. and This refers to the measured pitch angular velocity of the aircraft. This is the measured value of the yaw rate on the aircraft.

4. The robust control method for aircraft based on angular acceleration feedback according to claim 3, characterized in that, The desired longitudinal dynamics design is based on a proportional-integral controller, with angular acceleration and pitch angular velocity as feedback variables. The desired angular acceleration is expressed as follows: ； in, For command acceleration, Normal acceleration, The proportional control parameter for pitch angle acceleration. The integral control parameter for pitch angle acceleration, These are the control parameters for normal acceleration. These are the control parameters for pitch angular velocity; The short-period modes of longitudinal motion are represented as follows: ； in, For The actual speed of the aircraft is expressed in units of... The pitch angle time constant, Let be the derivative of the aircraft's angle of attack. For gravitational acceleration, the short-period modes of the aircraft can be represented by a transfer function. For complex variables in the transfer function; Combining the above two equations, we can obtain the transfer function as follows; ； The characteristic equation of the closed-loop system is: In order to eliminate , for: ； In the above formula, It is the damping ratio of the short-period mode. It is the natural frequency of the short-period mode; Then, the initial values ​​of the flight quality parameters are obtained as follows: ； You can also get: ； The pitch rate for control commands can be written from the aircraft's longitudinal equations as: ； Combining the above three formulas, we have: ； Pole-zero cancellation is performed on the pitch angular velocity expression: ； In the formula, The desired pitch angle time constant selected according to flight quality; An expression containing s; for The pole-related expression obtained after transformation.

5. The robust control method for an aircraft based on angular acceleration feedback according to claim 4, characterized in that, Related to command path pre-filter, The control parameters derived from the actual overload are used in the command path pre-filter. For the control parameters derived from command overload in the command path pre-filter: 。

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