Effect adjusting control algorithm suitable for inner ring zero-static-error control

Abstract

The invention belongs to the technical field of aviation flight control, and particularly relates to an effect regulation control algorithm suitable for inner loop zero static error control, a longitudinal corresponding inertia dynamic regulation link is added, an input quantity is calculated based on a longitudinal effect regulation mechanism control law to obtain a longitudinal effect regulation mechanism control rate for forming transient suppression, and the transient suppression is realized. Calculating based on the control rate of the longitudinal regulation mechanism to obtain a longitudinal regulation calculation result; adding a transverse corresponding dynamic inertia adjustment link, calculating an input quantity based on a transverse adjustment mechanism control law to obtain a transverse adjustment mechanism control rate forming transient suppression, and calculating a transverse adjustment calculation result based on the transverse adjustment mechanism control rate; the method can adapt to the effect regulation control design of a static-error-free aircraft inner ring control system, the problem of transient transformation in flight mode transformation can be well solved, it is guaranteed that an aircraft has a stable flight state, and the algorithm has the advantages of being simple in structure, high in expandability and capable of having a very high practical function.

Description

Technical Field

[0001] This application belongs to the field of aviation flight control technology, and specifically relates to an effect-adjusting control algorithm that adapts to inner-loop zero steady-state error control. Background Technology

[0002] Automatic flight control (AFC) is designed to reduce pilot fatigue during long-haul flights and can replace most of the pilot's operational functions. Modern aircraft typically use a zero-static-error fly-by-wire flight control system in their inner loop, while the AFC's control loop, located outside the inner loop, usually consists of pitch, roll, and yaw loops formed by attitude and angular velocity feedback. These three loops divide the pilot's basic maneuvers into multiple modes of aircraft angular and center-of-gravity motion, enabling the aircraft to be controlled according to the pilot's predetermined flight attitude and trajectory.

[0003] However, in these mode switching controls, during the transition between two modes, transient transitions may occur between modes, leading to aircraft vibration. This is because during automatic flight control, changes in the aircraft trim state can cause shifts in stick displacement. After the mode transition, the shift will deviate from the initial value of the other mode, resulting in transients.

[0004] Therefore, how to eliminate transient values ​​is a problem that needs to be solved. Summary of the Invention

[0005] The purpose of this application is to provide an adjustment control algorithm that adapts to the inner loop zero steady-state error control, so as to solve the problem of transients that easily occur when the aircraft trim state changes.

[0006] The technical solution of this application is: a tuning control algorithm adapted to inner-loop zero-steady-error control, comprising:

[0007] The difference between the longitudinal control quantity calculated by the automatic flight control system and the introduced overload feedback is used as the input quantity for solving the longitudinal tuning mechanism control law. A longitudinal inertial dynamic adjustment link is added, and the longitudinal tuning mechanism control law control law control law control law that forms transient suppression is obtained based on the input quantity. The longitudinal tuning solution result is calculated based on the longitudinal tuning mechanism control law.

[0008] The difference between the lateral control quantity calculated by the automatic flight control system and the introduced roll rate feedback is used as the input quantity for the lateral adjustment mechanism control law solution. A corresponding lateral dynamic inertial adjustment element is added, and the lateral adjustment mechanism control law control rate that forms transient suppression is obtained based on the lateral adjustment mechanism control law solution input quantity. The lateral adjustment solution result is obtained based on the lateral adjustment mechanism control rate.

[0009] Preferably, the longitudinal tuning solution is subjected to hysteresis transformation to obtain the automatic tuning output signal of the aircraft drive mechanism under different flight modes; the lateral tuning solution is subjected to hysteresis transformation to obtain the automatic tuning output signal of the aircraft drive mechanism under different flight modes.

[0010] Preferably, the control rate of the longitudinal modulation mechanism that forms transient suppression is:

[0011] ;

[0012] Where K1 is the longitudinal control variable of the automatic flight control system. The gain of the branch, K2 is the gain of the longitudinal overload ny branch, Given the time constant of the inertial element, the longitudinal adjustment solution is obtained. .

[0013] Preferably, when a longitudinally corresponding inertial dynamic adjustment element is added, the time constant of the inertial element is... The longitudinal parameters are adjusted according to the control loop; when the aircraft is controlled to pitch down, the on and off threshold parameters on1 and off1 are adjusted using the first quadrant hysteresis loop; when the aircraft is controlled to pitch up, the on and off threshold parameters on2 and off2 are adjusted using the fourth quadrant hysteresis loop; ultimately, the intermittent adjustment function is achieved.

[0014] Preferably, the control law formula for the lateral adjustment mechanism is:

[0015] ;

[0016] Where K3 is the longitudinal control variable of the automatic flight control system. Branch gain, K4 is for lateral overload Branch gain, Given the time constant of the inertial element, the transverse tuning solution is obtained. .

[0017] Preferably, when a corresponding lateral dynamic inertial adjustment element is added, the time constant of the inertial element is... The parameters are adjusted based on the control loop; the hysteresis loop parameters are calculated from the lateral adjustment results. When controlling the aircraft to roll to the left, the on and off threshold parameters on3 and off3 are adjusted using the first quadrant hysteresis loop; when controlling the aircraft to roll to the right, the on and off threshold parameters on4 and off4 are adjusted using the fourth quadrant hysteresis loop; ultimately, the intermittent adjustment function is achieved.

[0018] The tuning control algorithm adapted to inner-loop zero steady-state error control in this application can be adapted to the tuning control design of inner-loop control systems for zero steady-state error aircraft. It can effectively eliminate the problem of transient changes in flight mode transitions, ensuring that the aircraft has a stable flight state. Moreover, the algorithm has a simple structure, strong scalability, and strong practical functions, and has broad engineering application prospects in various types of aircraft. Attached Figure Description

[0019] To more clearly illustrate the technical solutions provided in this application, the accompanying drawings will be briefly described below. Obviously, the drawings described below are merely some embodiments of this application.

[0020] Figure 1 This is a block diagram of the vertical performance control structure of this application;

[0021] Figure 2 This is a block diagram of the horizontal adjustment control structure of this application. Detailed Implementation

[0022] 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.

[0023] A tuning control algorithm adapted to inner-loop zero steady-state error control, such as Figures 1-2 It includes the following steps:

[0024] Step S100: The difference between the longitudinal control quantity calculated by the automatic flight control system and the introduced overload feedback is used as the input quantity for the longitudinal adjustment mechanism control law solution; a longitudinal inertial dynamic adjustment link is added, and the longitudinal adjustment mechanism control law control rate that forms transient suppression is obtained based on the longitudinal adjustment mechanism control law solution input quantity, and the longitudinal adjustment solution result is calculated based on the longitudinal adjustment mechanism control rate.

[0025] Hysteresis transformation is performed on the longitudinal tuning solution to obtain the automatic tuning output signal of the aircraft drive mechanism under different flight modes.

[0026] Preferably, the control rate of the longitudinal modulation mechanism that forms transient suppression is:

[0027] ;

[0028] Where K1 is the longitudinal control variable of the automatic flight control system. The gain of the branch, K2 is the gain of the longitudinal overload ny branch, Given the time constant of the inertial element, the longitudinal adjustment solution is obtained. .

[0029] When a longitudinally corresponding inertial dynamic adjustment element is added, the time constant of the inertial element... The longitudinal parameters are adjusted according to the control loop; when the aircraft is controlled to pitch down, the on and off threshold parameters on1 and off1 are adjusted using the first quadrant hysteresis loop; when the aircraft is controlled to pitch up, the on and off threshold parameters on2 and off2 are adjusted using the fourth quadrant hysteresis loop; finally, the intermittent adjustment function is realized, and the aircraft drive mechanism is automatically adjusted under different flight modes.

[0030] The results of parameter adjustments are shown in Table 1:

[0031] Table 1. Gains and coefficients in the zero-steady-error inner-loop longitudinal tuning control method

[0032]

[0033] Step S200: The difference between the lateral control quantity calculated by the automatic flight control system and the introduced roll rate feedback is used as the input quantity for the lateral adjustment mechanism control law solution; a corresponding lateral dynamic inertial adjustment link is added; the lateral adjustment mechanism control law solution input quantity is used to obtain the lateral adjustment mechanism control law that forms transient suppression; and the lateral adjustment solution result is calculated based on the lateral adjustment mechanism control law solution input quantity.

[0034] Hysteresis transformation is performed on the lateral tuning solution to obtain the automatic tuning output signal of the aircraft drive mechanism under different flight modes.

[0035] Preferably, the control law formula for the lateral adjustment mechanism is:

[0036] ;

[0037] Where K3 is the longitudinal control variable of the automatic flight control system. Branch gain, K4 is for lateral overload Branch gain, Given the time constant of the inertial element, the transverse tuning solution is obtained. .

[0038] When a corresponding dynamic inertial adjustment element is added laterally, the time constant of the inertial element... The parameters are adjusted based on the control loop; the hysteresis loop parameters are calculated from the lateral adjustment results. When controlling the aircraft to roll to the left, the on and off threshold parameters on3 and off3 are adjusted using the first quadrant hysteresis loop; when controlling the aircraft to roll to the right, the on and off threshold parameters on4 and off4 are adjusted using the fourth quadrant hysteresis loop; ultimately, the intermittent adjustment function is achieved, and the aircraft drive mechanism automatically adjusts under different flight modes.

[0039] The results of parameter adjustments are shown in Table 2:

[0040] Table 2. Gains and coefficients in the zero-steady-error inner-loop longitudinal tuning control method.

[0041] In summary, this application has the following advantages:

[0042] This algorithm is adaptable to the tuning control design of the inner loop control system of aircraft with zero static error. It can effectively eliminate the problem of transient changes in flight mode transitions, ensuring that the aircraft has a stable flight state. Moreover, the algorithm has a simple structure, strong scalability, and strong practical functions, and has broad engineering application prospects in various types of aircraft.

[0043] Finally, it should be noted that the accompanying drawings of the embodiments disclosed in this invention only involve the structures involved in the embodiments disclosed in this invention. Other structures can refer to the general design. In the absence of conflict, the same embodiment and different embodiments of this invention can be combined with each other.

[0044] In conclusion, the above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A droop control algorithm for adaptive inner loop zero static error control, characterized in that, Comprise: The longitudinal control amount calculated by the automatic flight control system and the introduced overload feedback are subtracted as the input quantity of the longitudinal servo control law calculation; the corresponding inertia dynamic adjustment link in the longitudinal direction is added, the longitudinal servo control rate forming transient suppression is obtained based on the input quantity of the longitudinal servo control law calculation, and the longitudinal servo calculation result is calculated based on the longitudinal servo control rate; The lateral control amount calculated by the automatic flight control system and the introduced roll angular velocity feedback are subtracted as the input quantity of the lateral servo control law calculation; the corresponding dynamic inertia adjustment link in the lateral direction is added, the lateral servo control rate forming transient suppression is obtained based on the input quantity of the lateral servo control law calculation, and the lateral servo calculation result is calculated based on the lateral servo control rate.

2. The droop control algorithm for adaptive inner loop deadbeat control as claimed in claim 1 wherein: The longitudinal servo calculation result is hysteresis converted to obtain the automatic servo output signal of the aircraft driving mechanism under different flight modes; the lateral servo calculation result is hysteresis converted to obtain the automatic servo output signal of the aircraft driving mechanism under different flight modes.

3. The droop control algorithm for adaptive inner loop deadbeat control as claimed in claim 2 wherein: The longitudinal servo control rate forming transient suppression is: ； where K1 is the longitudinal control quantity of the automatic flight control system K2 is the gain of the longitudinal overload ny branch, is the inertial link time constant, and the longitudinal control solution is obtained .

4. The droop control algorithm for adaptive inner loop deadbeat control as claimed in claim 3 wherein, When the longitudinal corresponding inertia dynamic adjustment link is added, the inertia link time constant According to the control link, longitudinal parameters are adjusted; when the aircraft is pitching down, the first quadrant hysteresis link is used to adjust the on and off threshold parameters on1 and off1; when the aircraft is pitching up, the fourth quadrant hysteresis link is used to adjust the on and off threshold parameters on2 and off2; and finally the intermittent efficiency function is realized.

5. The droop control algorithm for adaptive inner loop deadbeat control as recited in claim 2, wherein, The lateral servo control law formula is: ； where K3 is the longitudinal control of the automatic flight control system Gain of the branch, K4 is the lateral overload Gain of the branch, is the inertial link time constant, and the lateral control solution is obtained .

6. The droop control algorithm for adaptive inner loop deadbeat control as recited in claim 5, wherein: When the corresponding dynamic inertia adjustment link is added in the lateral direction, the inertia link time constant The parameters of the hysteresis link are adjusted according to the control link; when the airplane is controlled to roll to the left, the first quadrant hysteresis link is used to adjust the on and off threshold parameters on3 and off3; when the airplane is controlled to roll to the right, the fourth quadrant hysteresis link is used to adjust the on and off threshold parameters on4 and off4; and finally the intermittent control function is realized.

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