Aircraft automatic landing PID longitudinal guidance law loop design method

Abstract

The invention belongs to the technical field of aviation aircraft control, and particularly relates to an aircraft automatic landing PID longitudinal guidance law loop design method, which comprises the following steps: calculating a second-order system and a third-order transfer function in a vertical rate guidance mode; a third-order transfer function is equivalent to a PID controller with feedforward filtering, parameters in the transfer function are set, a first-order filter, a second-order expected reference model and a forward PD compensator of the PID controller are calculated, and the PID controller is established. And substituting the PID controller back to the original system, adjusting various gains of the original system, and carrying out longitudinal guide control on automatic landing of the aircraft. According to the method, the mapping relation between the filter, the expected transfer function, the regulator parameter and the longitudinal guidance law loop PID controller parameter in the anti-interference structure can be established, a large number of attempt processes in an experience-based parameter regulation scheme are avoided, and the efficiency of control design is remarkably improved.

Description

Technical Field

[0001] This application belongs to the technical field of aircraft control, and particularly relates to a design method for the PID longitudinal guidance law loop of aircraft automatic landing. Background Art

[0002] In the aircraft automatic landing control system, the function of the guidance law loop is to form a guidance command according to a certain rule based on the difference between the aircraft altitude measured by sensors such as radar and the specified glide altitude, and input it into the inner loop such as the automatic flight control system. The general implementation method of the guidance law loop is to convert the altitude error during landing into a vertical speed command through a controller and use it as the input of the automatic flight control system.

[0003] The form of the guidance law loop controller is generally a PID (Proportional-Integral-Derivative) form controller. This type of controller does not depend on the specific form of the model for single-input single-output and has good control effects. However, determining each control gain depends on engineering experience, and there is a lot of work in adjusting parameters and trial and error.

[0004] Control strategies based on disturbance observers, such as active disturbance rejection control, can also be applied to most systems, do not overly rely on the accurate model of the controlled object, and have certain robustness.

[0005] Therefore, how to design a more effective control strategy for aircraft automatic landing is a problem that needs to be solved. Summary of the Invention

[0006] The purpose of this application is to provide a design method for the PID longitudinal guidance law loop of aircraft automatic landing to solve the problem of more trial and error work in the existing control strategy of using PID to control aircraft automatic landing.

[0007] The technical solution of this application is: a design method for the PID longitudinal guidance law loop of aircraft automatic landing, including: Connect the longitudinal linearized subsystem of the aircraft to the inner control loop, ignore the zeros and poles far from the imaginary axis in the inner control loop, and calculate the second-order system and the third-order transfer function in the vertical speed guidance mode; Equivalent the third-order transfer function to a PID controller with feedforward filtering, and then based on the control method of the disturbance observer, improve the control loop of the PID controller with feedforward filtering to obtain an equivalent control loop. Then, based on the equivalent control loop, calculate the first-order filter, the second-order expected reference model, and the forward PD compensator of the PID controller to establish the PID controller; Determine the values of each gain of the PID controller, substitute the PID controller back into the original system, adjust the gains of the original system, and perform longitudinal guidance control on aircraft automatic landing.

[0008] Preferably, the vertical rate is commanded to the aircraft sink rate in the vertical rate guidance mode The transfer function of the channel is configured as a second-order system The second-order system The expression is; ; Wherein, Is the vertical rate command, calculated by the controller of the guidance law loop; Is the rate of change of the aircraft's altitude in the vertical sea surface direction, reflecting the longitudinal descent process of the aircraft. 、 And Are all numerator polynomial coefficients; 、 、 Are all denominator polynomial coefficients, 、 、 、 Are all Laplace operators; N is the numerator polynomial, Represents the number of terms, and D is the denominator polynomial.

[0009] Preferably, in the second-order system, the actual disturbance altitude h of the aircraft is used as the output to obtain the third-order transfer function of the aircraft sink rate Channel, which is: .

[0010] Preferably, the first-order filter Is designed as: ; The second-order desired reference model Is designed as: ; The forward PD compensator Is designed as: ; In the formula, s is the Laplace operator, Is a fixed constant, Is a fixed constant, Is the time constant, Is the damping ratio, Is the frequency.

[0011] Preferably, the fixed constant 、Time constant 、Damping ratio 、Frequency The conditions satisfied are: In the second-order desired reference model is a Hurwitz polynomial. In the forward PD compensator In the first-order filter And let the polynomial be: ; is a Hurwitz polynomial.

[0012] Preferably, the PID controller is designed as: ; In the formula, is the integral gain, is the proportional gain, is the derivative gain; is the second derivative gain; And: .

[0013] The design method of the PID longitudinal guidance law loop for automatic landing of an aircraft in this application can establish the mapping relationship between the filter, the desired transfer function, the regulator parameters in the disturbance rejection structure and the PID controller parameters of the longitudinal guidance law loop, avoiding a large number of trial processes in the empirical tuning parameter scheme and significantly improving the efficiency of control design. Brief Description of the Drawings

[0014] In order to more clearly illustrate the technical solutions provided in this application, the drawings will be briefly introduced below. Obviously, the drawings described below are only some embodiments of this application.

[0015] Figure 1 is the overall flow schematic diagram of this application; Figure 2 is the structural block diagram of this application based on the disturbance observer form; Figure 3 is the equivalent PID control structural block diagram of the guidance law loop of this application. Detailed Description of the Embodiments

[0016] The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are only some of the embodiments of the present invention, rather than all of them. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0017] A design method of the PID longitudinal guidance law loop for automatic landing of an aircraft, as Figure 1 , includes the following steps: Step S100, connect the longitudinal linearization subsystem of the aircraft to the inner control loop, ignore the zeros and poles far from the imaginary axis in the inner control loop, and calculate the second-order system and the third-order transfer function in the vertical rate guidance mode; Preferably, in the vertical rate guidance mode, the transfer function of the vertical rate command to the aircraft sink rate channel is configured as a second-order system , and the expression of the second-order system is; ; where, is the vertical rate command, which is calculated by the controller of the guidance law loop; is the rate of change of the aircraft's altitude in the vertical sea surface direction, which reflects the longitudinal descent process of the aircraft. , and are all the coefficients of the numerator polynomial; , , are all the coefficients of the denominator polynomial, , , , are all Laplace operators. N is the numerator polynomial and D is the denominator polynomial.

[0018] The transfer function is obtained as a second-order system through loop order reduction or frequency domain equivalent configuration method. Since there are a large number of filter and actuator links with zeros and poles far from the imaginary axis in the original model, the low-order model obtained by ignoring such links has a high degree of approximation.

[0019] According to the control requirements of the inner control loop for the aircraft sink rate channel, the second-order system can ensure that both the poles and zeros of the transfer function are in the left half complex open plane.

[0020] Taking the actual disturbance altitude h of the aircraft as the output in the second-order system, the third-order transfer function of the aircraft sink rate channel is: .

[0021] Step S200, equivalent the third-order transfer function to a PID controller with feedforward filtering, and then, based on the control method of the disturbance observer, improve the control loop of the PID controller with feedforward filtering, such as Figure 2 , to obtain an equivalent control loop, such as Figure 3 , and then calculate the first-order filter, the second-order desired reference model and the forward PD compensator of the PID controller based on the equivalent control loop to establish the PID controller.

[0022] Preferably, the first-order filter Designed as: ; Second-order expected reference model Designed as: ; Forward PD compensator Designed as: ; In the formula, s is the Laplace operator, is a fixed constant, is a fixed constant, is the time constant, is the damping ratio, is the frequency. The fixed constant and the time constant are both set according to experience.

[0023] is a fixed constant, is the time constant, is the damping ratio, The condition that the frequency satisfies is: In the second-order expected reference model is a Hurwitz polynomial. In the forward PD compensator , in the first-order filter And let the polynomial: ; is a Hurwitz polynomial.

[0024] Preferably, the PID controller is designed as: ; In the formula, is the integral gain, is the proportional gain, is the derivative gain; is the second derivative gain.

[0025] .

[0026] Step S300, determine the values of the gains of the PID controller, including the frequency , the damping ratio , the integral gain , the proportional gain , the derivative gain and the second derivative gain , substitute the PID controller back into the original system, adjust the gains of the original system, and perform longitudinal guidance control for the automatic landing of the aircraft. Make the aircraft have a good tracking effect on the altitude glide slope command under the guidance law loop.

[0027] Preferably, the specific values of the gains of the PID controller are determined as follows: According to the stability requirement and the rapidity requirement of the system response, the time constant is selected to ensure stability and rapidity. According to the coefficient in the transfer function, the value of the second-order differential gain is determined.

[0028] According to the bandwidth of the second-order desired reference model and the response satisfying the desired aircraft altitude guidance response, the frequency and damping ratio in the second-order desired reference model are selected.

[0029] According to the requirement that the closed-loop response of the low-order fitting system is satisfied, combined with the damping ratio the integral gain , proportional gain and differential gain are adjusted as a whole.

[0030] In summary, the method for designing the PID longitudinal guidance law loop of the automatic landing of an aircraft based on an equivalent disturbance rejection structure according to the present application can establish the mapping relationship between the filter, desired transfer function, regulator parameters in the disturbance rejection structure and the PID controller parameters of the longitudinal guidance law loop, avoiding a large number of trial processes in the empirical tuning parameter scheme and significantly improving the efficiency of control design.

[0031] Finally, it should be noted that: in the accompanying drawings of the disclosed embodiments of the present invention, only the structures related to the disclosed embodiments are involved. For other structures, reference may be made to the general design. Without conflict, the same embodiment and different embodiments of the present invention can be combined with each other; Finally: The above are only the preferred embodiments of the present invention and are not used to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principle of the present invention shall be included in the protection scope of the present invention.

Claims

1. A design method for an aircraft automatic landing PID longitudinal guidance law loop, characterized in that: include: Connect the aircraft longitudinal linearization subsystem to the inner control loop, ignore the zero poles far from the imaginary axis in the inner control loop, and calculate the second-order system and third-order transfer functions in the vertical rate guidance mode. The third-order transfer function is equivalent to a PID controller with feedforward filtering, and then based on the control method of the disturbance observer, the control loop of the PID controller with feedforward filtering is improved to obtain an equivalent control loop. Then, based on the equivalent control loop, the first-order filter, the second-order desired reference model and the forward PD compensator of the PID controller are calculated to establish the PID controller. Determine the values ​​of each gain of the PID controller, substitute the PID controller back into the original system, adjust the various gains of the original system, and perform longitudinal guidance control for the automatic landing of the aircraft.

2. The method for designing an aircraft automatic landing PID longitudinal guidance law loop according to claim 1, characterized in that: In vertical rate guidance mode, the vertical rate command is set to the aircraft sink rate. The transfer function of the channel is configured as a second-order system , second-order system The expression is; ； in, is the vertical rate command, which is calculated by the controller of the guidance law loop; It is the rate of change of the aircraft's altitude in the direction perpendicular to the sea surface, reflecting the aircraft's longitudinal descent process; , and are the coefficients of the numerator polynomial; , , are the coefficients of the denominator polynomial, , , , are all Laplace operators; N is the numerator polynomial, represents the number of terms, and D is the denominator polynomial.

3. The method for designing an aircraft automatic landing PID longitudinal guidance law loop according to claim 2, characterized in that: In the second-order system, the actual disturbance height h of the aircraft is used as the output to obtain the aircraft sinking rate The third-order transfer function of the channel is: 。 4. The method for designing an aircraft automatic landing PID longitudinal guidance law loop as claimed in claim 2, characterized in that: The first-order filter Designed for: ； The second-order expected reference model Designed for: ； The forward PD compensator Designed for: ； Where s is the Laplace operator, is a fixed constant, is a fixed constant, is the time constant, is the damping ratio, is the frequency.

5. The method for designing an aircraft automatic landing PID longitudinal guidance law loop according to claim 4, characterized in that: The fixed constant , time constant , Damping ratio ,frequency The conditions met are: Second-order expected reference model is the Hurwitz polynomial, in the forward PD compensator , in a first-order filter And let the polynomial be: ； is the Hurwitz polynomial.

6. The method for designing an aircraft automatic landing PID longitudinal guidance law loop as claimed in claim 4, characterized in that: The PID controller is designed as: ； In the formula, is the integral gain, is the proportional gain, is the differential gain; is the second derivative gain; and: 。

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