CN117930632A - High-reliability safe flight control method for enhancing stable reserve of system - Google Patents
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Abstract
The invention provides a high-reliability safe flight control method for enhancing stable reserve of a system, which utilizes the effect of angular acceleration signals on the system by directly reacting to a moment of stress, and based on the characteristic that the system can enhance the robustness of the system to disturbance moment and load fluctuation when the angular acceleration signals are fed back into a servo system, the angular acceleration signals are applied to the flight control system to further improve the stability and the robustness of an aircraft, the stable reserve of the flight control system is increased, the problem that the flight performance of the carrier is influenced by flow separation caused by a bait cabin and a towing line after the carrier releases towing type lure bait tank is solved, and a better control effect is obtained.
Description
Technical Field
The invention belongs to the technical field of aircraft control methods, and particularly relates to a high-reliability safe flight control method for enhancing stable reserve of a system.
Background
The towing type bait cabin is configured with the carrier through the towing line, is similar to the flying speed and track of the carrier, can interfere enemy detection signals, and improves the flying safety of the carrier. Because the lure bait tank flies by means of the towing line, the following motor action of the carrier has hysteresis, the condition of the carrier tail flow wind field is complex, and the towing type bait cabin easily threatens the flight safety of the carrier. Therefore, there is a need to provide a highly reliable and safe flight control method capable of improving the flight safety of a carrier and accurately controlling the flight under the influence of a complex wind field caused by the existence of a towed bait cabin.
Specifically, in the attitude angle control system of the conventional flight control system, a control strategy is generally built according to a time scale separation principle. The cascade control strategy based on attitude angle and attitude angular speed control is difficult to realize good control effect under the conditions of model uncertainty, parameter uncertainty, gravity center perturbation and the like, namely the stable reserve of the system is insufficient. The classical design method depends on an accurate model, can not realize a better control effect when the model is inaccurate, and can not quickly return to a stable state when the system is disturbed. When the carrier releases the drag type induced bait tank, the complex wind field caused by the flow separation caused by the bait cabin and the drag line affects the flight performance of the carrier, in addition, the flow separation can cause moment interference on the pitching axis of the carrier and affect the deflection of the control surface, so that model uncertainty and parameter uncertainty are caused, and the flight performance of the carrier is also affected. Therefore, it is difficult for the conventional flight control system to achieve precise and stable safety control under such circumstances.
Disclosure of Invention
Aiming at the problems existing in the prior art, the invention provides a high-reliability safe flight control method for enhancing the stable reserve of a system, which utilizes the effect of an angular acceleration signal on a system and can directly react to a moment, based on the characteristic that the system can enhance the robustness of the system to disturbance moment and load fluctuation when the angular acceleration signal is fed back into a servo system, applies the angular acceleration signal to the flight control system to further improve the stability and the robustness of an aircraft, increases the stable reserve of the flight control system, overcomes the problem that the flight performance of the carrier is influenced by the flow separation caused by a bait cabin and a towing line after the carrier releases towing type lure bait tank, and obtains a better control effect.
The technical scheme of the invention is as follows:
A high reliability safe flight control method for enhancing system stability reserve comprising the steps of:
Step 1: establishing a carrier aircraft model with a drag type attractant bait tank, dividing a posture angle control system of the carrier aircraft with the drag type attractant bait tank into a posture angle link and a posture angle speed link according to a time mark separation principle, and designing the aircraft posture angle control system based on a proportional-integral-derivative control method;
Step 2: based on an angular acceleration feedback principle, adding an attitude angular acceleration link after the attitude angular speed link of the aircraft attitude angular control system, and designing an attitude angular acceleration controller based on a proportional-integral-derivative control method to obtain an aircraft attitude angular control system with enhanced stable reserve; and the estimated value of the attitude angle acceleration signal fed back in the attitude angle acceleration link is obtained based on a complementary filtering principle.
Further, in step 2, the process of obtaining the estimated value of the attitude angular acceleration signal based on the complementary filtering principle is as follows:
And obtaining an initial angular acceleration signal from an attitude angular velocity signal obtained by an airborne sensor through a differential link, integrating the initial angular acceleration signal, and making a difference with the attitude angular velocity signal, obtaining a compensation control quantity after the difference value passes through a proportional-integral controller, and compensating the compensation control quantity into the initial angular acceleration signal to realize a feedforward process and obtain an estimated value of the attitude angular acceleration signal.
Further, in step 2, the expression for obtaining the estimated value of the attitude angular acceleration signal based on the complementary filtering principle is:
wherein, For the attitude angular velocity signal measurements obtained by the on-board sensor,For the derivative value of the attitude angular velocity signal obtained by the on-board sensor,For the estimated value of the attitude angle acceleration signal,For the purpose of the proportional control parameter,In order to integrate the control parameters,Is a laplace operator.
Further, the attitude angle control system of the aircraft with drag-type attraction bait tank designed in the step1 comprises:
pitch angle control loop: the pitch angle and pitch angle speed controller in the loop is designed as follows:
wherein, As a desired value of the pitch angle rate,As a measure of the pitch angle rate,Is the desired value of the pitch angle,As a measure of the pitch angle,Is a proportional control parameter in a pitch angle proportional integral controller,Is an integral control parameter in a pitch angle proportional integral controller,Is a proportional control parameter in a pitch rate proportional integral controller,For the integral control parameter in the pitch rate proportional integral controller,For the deflection angle of the elevator,The trim angle for the elevator.
Roll angle control circuit: the roll angle and roll angle speed controller in the loop is designed as follows:
wherein, As a desired value of the roll angle speed,As a measure of the roll angle velocity,As a desired value of the roll angle,As a measure of the roll angle,As a proportional control parameter in the roll angle proportional integral controller,As an integral control parameter in the roll angle proportional integral controller,For the proportional control parameter in the roll angle speed proportional integral controller,For the integral control parameter in the roll angle speed proportional integral controller,For the deflection angle of the aileron,The trim angle for the aileron.
Yaw angle control loop: the yaw angle and yaw rate controller in the loop is designed to:
wherein, As the desired value of the yaw rate,As a measure of the yaw rate,As a desired value of the yaw angle,As a measure of the yaw angle,For the proportional control parameter in the yaw proportional integral controller,For the integral control parameter in the yaw angle proportional-integral controller,For the proportional control parameter in the yaw rate proportional integral controller,For the integral control parameter in the yaw rate proportional integral controller,For the deflection angle of the rudder,The trim angle for the rudder.
Further, in step 2, the aircraft attitude angle control system for enhancing the stable reserve includes:
Pitch angle control loop: the pitch angle, pitch angle speed and pitch angle acceleration controller in the loop are designed as follows:
wherein, Is the desired value of the pitch angle acceleration signal,Is an estimate of the pitch angle acceleration signal,For proportional control parameters in the pitch angle acceleration proportional integral controller,And integrating control parameters of a proportional-integral controller for the pitch angle acceleration.
Roll angle control circuit: the roll angle, the roll angle speed and the roll angle acceleration controller in the loop are designed as follows:
wherein, For a desired value of the roll angle acceleration signal,As an estimate of the roll angle acceleration signal,For the proportional control parameter in the roll angle acceleration proportional integral controller,And (3) integrating control parameters in a proportional-integral controller for the rolling angle acceleration.
Yaw angle control loop: yaw angle, yaw rate and yaw acceleration controllers in the circuit are designed as follows:
wherein, For a desired value of the yaw acceleration signal,As an estimate of the yaw acceleration signal,For proportional control parameters in the yaw acceleration proportional integral controller,And integrating control parameters of a proportional-integral controller for yaw angular acceleration.
Further, in step 2, sensor noise is introduced into the attitude angle measurement value and the attitude angular speed measurement value in the aircraft attitude angle control system for enhancing the stable reserve.
Furthermore, the noise of the attitude angle sensor adopts a Gaussian white noise model, and the time delay of the angle sensor is introduced for 40e-3 seconds after the noise is added; the noise of the attitude angular velocity sensor adopts a Gaussian white noise model, the speed limit of 600deg/s is carried out on the attitude angular velocity, and the time delay of the angular velocity sensor of 15e-3 seconds is introduced after the noise is added.
Advantageous effects
The high-reliability safe flight control method for enhancing the stable reserve of the system, provided by the invention, is based on the characteristic that the robustness of the system can be enhanced by the feedback control of the angular acceleration signal, and the angular acceleration signal is introduced into the flight control system to obtain a better control effect, and has the specific advantages that:
(1) Compared with the classical control method, the flight control method introducing the angular acceleration signal has the advantages that under the condition that the time domain performance is similar, the amplitude margin is effectively improved, the stability of the flight control system of the aircraft can be effectively enhanced by the aid of the stable reserve improvement, and excellent control effects are achieved under the conditions of center of gravity perturbation, uncertain models and uncertain parameters;
(2) The novel estimation method of the angular acceleration signal based on the complementary filtering principle is provided, and under the condition of considering time delay and noise of the angle sensor and the angular velocity sensor in actual engineering, the angular acceleration signal with excellent quality can be obtained, and the engineering usability of the signal is ensured.
Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention.
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The foregoing and/or additional aspects and advantages of the invention will become apparent and may be better understood from the following description of embodiments taken in conjunction with the accompanying drawings in which:
FIG. 1 is a schematic diagram of the attitude angle control law architecture of the present invention;
FIG. 2 is a schematic diagram of a method for estimating an angular acceleration signal based on the principle of complementary filtering according to the present invention;
FIG. 3 is a graph showing the estimation effect of the method for estimating the angular acceleration signals when no sensor noise is introduced in the embodiment of the present invention;
FIG. 4 is a graph showing the estimation effect of the method for estimating each angular acceleration signal when sensor noise is introduced in the embodiment of the present invention;
FIG. 5 is a graph comparing longitudinal tracking performance of the control method in the embodiment of the present invention with that of the classical control method;
FIG. 6 is a graph of amplitude versus frequency characteristics of a longitudinal control loop of a flight control system based on a classical control method with an amplitude margin of 27.8dB in an embodiment of the invention;
FIG. 7 is a graph of amplitude versus frequency characteristics of a longitudinal control loop of a safety flight control system for enhancing system stability reserve in an embodiment of the present invention with an amplitude margin of 54.3dB;
FIG. 8 is a graph comparing the longitudinal tracking performance of the control method in the embodiment of the invention with that of the classical control method as the static margin is reduced;
FIG. 9 is a graph comparing the longitudinal tracking performance of the control method in the example of the present invention with that of the classical control method when the steering efficiency is reduced by 30%;
FIG. 10 is a graph comparing the longitudinal tracking performance of the control method in the example of the present invention with that of the classical control method when the steering efficiency is increased by 30%;
FIG. 11 is a graph comparing the longitudinal tracking performance of the control method in the embodiment of the present invention with that of the classical control method when the torque coefficient is reduced by 30%;
Fig. 12 is a graph comparing the longitudinal tracking performance of the control method in the embodiment of the present invention with that of the classical control method when the torque coefficient is increased by 30%.
Detailed Description
The following detailed description of embodiments of the invention is exemplary and intended to be illustrative of the invention and not to be construed as limiting the invention.
The high-reliability safe flight control method for enhancing the stable reserve of the system, provided by the invention, utilizes the characteristic that the angular acceleration signal directly reflects the action effect of the interference moment on the system, applies the angular acceleration signal to the flight control system of the towed bait cabin carrier to further improve the stability and robustness of the carrier, and increases the stable reserve of the carrier, so that the carrier does not reduce the flight performance due to the external interference caused by the bait cabin after the bait cabin is released bait tank.
The method specifically comprises the following steps:
Step 1: firstly, a model of a carrier aircraft with a drag type attractant bait tank is established, a posture angle control system of the carrier aircraft with the drag type attractant bait tank is divided into a posture angle link and a posture angle speed link according to a time mark separation principle, and a classical posture angle control system of the aircraft is designed based on a proportional-integral-derivative control method.
The three-axis attitude angle control system of the aircraft is designed based on a cascade control strategy:
The outer ring is an attitude angle control ring, and an expected attitude angle speed instruction is generated by utilizing an attitude angle tracking error through an attitude angle controller; the inner ring is an attitude angular speed control ring, and an aircraft control input signal is generated through an attitude angular speed controller by utilizing an attitude angular speed instruction generated by the outer ring.
Specifically:
In a pitch control loop of an aircraft, the pitch controller is:
wherein, As a desired value of the pitch angle rate,Is the desired value of the pitch angle,As a measure of the pitch angle,Is a proportional control parameter in a pitch angle proportional integral controller,Is an integral control parameter in a pitch angle proportional integral controller.
The pitch angle speed controller is:
wherein, For the deflection angle of the elevator,For the trim angle of the elevator,Is a proportional control parameter in a pitch rate proportional integral controller,For the integral control parameter in the pitch rate proportional integral controller,Is a measure of pitch angle rate.
In a roll angle control loop of an aircraft, a roll angle controller is:
wherein, As a desired value of the roll angle speed,As a desired value of the roll angle,As a measure of the roll angle,As a proportional control parameter in the roll angle proportional integral controller,Is an integral control parameter in the roll angle proportional integral controller.
The roll angle speed controller is:
wherein, For the deflection angle of the aileron,Is the trim angle of the aileron,For the proportional control parameter in the roll angle speed proportional integral controller,For the integral control parameter in the roll angle speed proportional integral controller,Is a measure of roll angle speed.
In a yaw angle control loop of an aircraft, a yaw angle controller is:
wherein, As the desired value of the yaw rate,As a desired value of the yaw angle,As a measure of the yaw angle,For the proportional control parameter in the yaw proportional integral controller,Is an integral control parameter in a yaw angle proportional integral controller.
The yaw rate controller is:
wherein, For the deflection angle of the rudder,Is the trim angle of the rudder,For the proportional control parameter in the yaw rate proportional integral controller,For the integral control parameter in the yaw rate proportional integral controller,Is a measure of yaw rate.
Step 2: based on the angular acceleration feedback principle, an attitude angular acceleration loop is added after an attitude angular speed link of an aircraft attitude angular control system, and an attitude angular acceleration controller is designed based on a proportional-integral-derivative control link.
And (2) after the step (1) of establishing the attitude angular speed link in the classical three-axis attitude angle control system of the aircraft, adding an attitude angular acceleration link, and designing an attitude angular acceleration controller based on the proportional-integral-derivative control link.
In the pitch control loop, the pitch rate controller is modified to:
wherein, Is the expected value of the pitch angle acceleration signal.
The pitch angle controller is unchanged, and the pitch angle acceleration controller is increased to be:
wherein, Is an estimate of the pitch angle acceleration signal,For proportional control parameters in the pitch angle acceleration proportional integral controller,And integrating control parameters of a proportional-integral controller for the pitch angle acceleration.
In the roll angle control loop, the roll angle speed controller is changed into:
wherein, Is the desired value of the roll angle acceleration signal.
The roll angle controller is unchanged, and the roll angle acceleration controller is added as follows:
wherein, As an estimate of the roll angle acceleration signal,For the proportional control parameter in the roll angle acceleration proportional integral controller,And (3) integrating control parameters in a proportional-integral controller for the rolling angle acceleration.
In the yaw angle control loop, a yaw angle speed controller is changed into:
wherein, Is the desired value of the yaw acceleration signal.
The yaw angle controller is unchanged, and the yaw angle acceleration controller is added as follows:
wherein, As an estimate of the yaw acceleration signal,For proportional control parameters in the yaw acceleration proportional integral controller,And integrating control parameters of a proportional-integral controller for yaw angular acceleration.
The estimated values, rather than the measured values, of the three attitude angular acceleration signals are mainly because no high-precision angular accelerometer can directly output the angular acceleration signals at present, so the current angular acceleration signals are always obtained according to a differential method or a Kalman filtering method based on the angular velocity signals, but the traditional method has large error and serious time delay when noise exists, and the invention adopts the method based on the complementary filtering principle to estimate the angular acceleration signals; and compared with classical angular acceleration estimation methods, namely a differential method and a Kalman filtering method.
First, an angular acceleration signal estimation method based on the complementary filtering principle is described. As shown in fig. 2, the angular velocity signal is first subjected to a differentiation step to obtain an angular acceleration signal, which is the same as that obtained by a conventional differentiation method, but the integral of the angular acceleration signal is then differentiated from the angular velocity signal and is subjected to a proportional-integral controller, and the control amount is compensated into the angular acceleration signal, thereby realizing a feedforward process.
The above procedure in the frequency domain can be written as:
wherein, For the attitude angular velocity signal measurements obtained by the on-board sensor,For the derivative value of the attitude angular velocity signal obtained by the on-board sensor,For the estimated value of the attitude angle acceleration signal,As a proportional control parameter in a proportional-integral controller,Is an integral control parameter in a proportional-integral controller,Is a laplace operator.
In actual flight, sensor noise also causes the performance of control law to be reduced, so that noise models of each attitude angle and attitude angular speed in a flight control system are respectively built, noise is introduced into the flight control system in a Gaussian white noise form, the average value of signal noise is set to be zero because the average value in a communication channel is generally zero, and the angle sensor noise modeling is respectively as follows:
Gaussian noise in roll angle sensor is noted as The one-dimensional probability density distribution of gaussian noise is:
Gaussian noise in pitch sensor is noted as The one-dimensional probability density distribution of gaussian noise is:
gaussian noise in yaw sensor is recorded as The one-dimensional probability density distribution of gaussian noise is:
The time delay of the angle sensor is introduced again after adding the noise, 40e-3 seconds.
The noise modeling of each angular velocity sensor is as follows:
Gaussian noise in a roll angle speed sensor is noted as The one-dimensional probability density distribution of gaussian noise is:
gaussian noise in pitch rate sensor is noted as The one-dimensional probability density distribution of gaussian noise is:
Gaussian noise in yaw rate sensor is recorded as The one-dimensional probability density distribution of gaussian noise is:
After adding noise, the upper and lower limits of the pitch, roll and yaw angular velocities are 600deg/s, and the time delay of the angular velocity sensor is 15e-3 seconds, in combination with the rate limitation of the actual angular velocity.
Based on the steps, the finally obtained control law architecture of the safety flight control system for enhancing the stable reserve of the system is shown in fig. 1.
The following comparative verification was performed for the above method:
Firstly, the proposed angular acceleration signal estimation method is compared with a differentiation method, a Kalman filtering method and an angular acceleration signal true value.
Fig. 2 is a schematic diagram of an angular acceleration signal estimation method based on the complementary filtering principle. FIG. 3 is a graph showing the estimation effect of the method for estimating the angular acceleration signals when no sensor noise is introduced in the embodiment of the present invention; the Kalman filtering response is obviously larger in time delay, the differential signal and the complementary filtering response are smaller in time delay and are closer to the true value, and the differential method is closest to the true value when no signal noise is introduced. Fig. 4 shows the estimation effect of each angular acceleration signal estimation method when sensor noise is introduced in the embodiment of the invention, and it can be obviously seen that the differential method response estimation noise is too large, the true value is almost submerged in the noise, the time delay of the kalman filter response is large, and the proposed model-independent angular acceleration signal estimation method, namely, the noise and the time delay level of the complementary filtering method are moderate. Table 1 shows delay time and signal to noise ratio of various methods corresponding to fig. 4, and it can be seen that the proposed angular acceleration signal estimation method based on the complementary filtering principle can better filter noise, obtain a signal with better quality, and meanwhile, the delay time is smaller.
Table 1 corresponds to the delay times and signal-to-noise ratios of the various methods of fig. 4, in which the complementary filtering method is even the angular acceleration signal estimation method based on the complementary filtering principle proposed in the present invention.
Table 1: delay time and noise ratio of each method
Secondly, the control method is used for longitudinal flight state simulation control of the aircraft, a cruise state is used as a working point representing flight characteristics, a time domain simulation experiment is carried out, frequency domain analysis is carried out, and control effects of two control loops using an angular velocity signal and an angular acceleration signal as feedback signals are compared. The comparison of the longitudinal tracking effect of the aircraft of the proposed control method and of the classical control method is shown in fig. 5; the system stability reserve when the aircraft is introduced into the proposed control method and the classical control method is shown in fig. 6 and 7, it can be seen that the introduction of the angular acceleration control link effectively enhances the stability reserve of the aircraft flight control system.
When the static margin of the test aircraft is reduced from 10% to 5%, the longitudinal tracking effect of the proposed control method and the classical control method is compared as shown in fig. 8, but the convergence speed of the proposed control method is high, the overshoot is small, and the stability of the flight control system can be effectively improved;
When the control efficiency of the test aircraft is biased, namely reduced by 30% and increased by 30%, longitudinal tracking effects of the control method and the classical control method are shown in figures 9 and 10, and the simulation result shows that the two control strategies can realize good attitude angle control, but the stability of the control method is better;
When the moment coefficient of the test aircraft is biased, namely, is reduced and increased by 30%, longitudinal tracking effects of the proposed control method and the classical control method are shown in fig. 11 and 12, and simulation results show that the two control methods can realize good attitude angle control, but the proposed control method has high convergence rate and more stable control effect.
In general, the flight control system introducing the angular acceleration signal control link can enhance the stable reserve of the flight control system, has a stable control effect, and has better reliability and safety when the aircraft faces gravity center perturbation, maneuvering efficiency deflection and flight moment coefficient deflection, namely, an aircraft model with parameter uncertainty.
Although embodiments of the present invention have been shown and described above, it will be understood that the above embodiments are illustrative and not to be construed as limiting the invention, and that variations, modifications, alternatives, and variations may be made in the above embodiments by those skilled in the art without departing from the spirit and principles of the invention.
Claims (7)
1. A high-reliability safe flight control method for enhancing stable reserve of a system is characterized in that: the method comprises the following steps:
Step 1: establishing a carrier aircraft model with a drag type attractant bait tank, dividing a posture angle control system of the carrier aircraft with the drag type attractant bait tank into a posture angle link and a posture angle speed link according to a time mark separation principle, and designing the aircraft posture angle control system based on a proportional-integral-derivative control method;
Step 2: based on an angular acceleration feedback principle, adding an attitude angular acceleration link after the attitude angular speed link of the aircraft attitude angular control system, and designing an attitude angular acceleration controller based on a proportional-integral-derivative control method to obtain an aircraft attitude angular control system with enhanced stable reserve; and the estimated value of the attitude angle acceleration signal fed back in the attitude angle acceleration link is obtained based on a complementary filtering principle.
2. A high reliability safe flight control method for enhancing system stability reserve as claimed in claim 1, wherein: in the step2, the process of obtaining the estimated value of the attitude angular acceleration signal based on the complementary filtering principle is as follows:
And obtaining an initial angular acceleration signal from an attitude angular velocity signal obtained by an airborne sensor through a differential link, integrating the initial angular acceleration signal, and making a difference with the attitude angular velocity signal, obtaining a compensation control quantity after the difference value passes through a proportional-integral controller, and compensating the compensation control quantity into the initial angular acceleration signal to realize a feedforward process and obtain an estimated value of the attitude angular acceleration signal.
3. A highly reliable and safe flight control method for enhancing system stability reserve according to claim 1 or 2, characterized in that: in the step 2, the expression for obtaining the estimated value of the attitude angular acceleration signal based on the complementary filtering principle is as follows:
wherein, For attitude angular velocity signal measurements obtained by an on-board sensor,/>Differential value of attitude angular velocity signal obtained for airborne sensor,/>For the estimated value of the attitude angle acceleration signal,/>Is a proportional control parameter,/>In order to integrate the control parameters,Is a laplace operator.
4. A high reliability safe flight control method for enhancing system stability reserve as claimed in claim 1, wherein: the attitude angle control system of the aircraft with drag-type attraction bait tank designed in the step 1 comprises the following steps:
pitch angle control loop: the pitch angle and pitch angle speed controller in the loop is designed as follows:
wherein, Is the expected value of pitch angle rate,/>Is a measurement of pitch angle rate,/>Is the expected value of pitch angle,/>Is a measurement of pitch angle,/>Is a proportional control parameter in a pitch angle proportional integral controller,/>Is an integral control parameter in a pitch angle proportional integral controller,/>Is a proportional control parameter in a pitch angle speed proportional integral controller,/>For integral control parameters in a pitch rate proportional integral controller,/>For the deflection angle of the elevator,/>Balancing angles for the elevators;
roll angle control circuit: the roll angle and roll angle speed controller in the loop is designed as follows:
wherein, Is the expected value of roll angle speed,/>As a measure of roll angle velocity,/>Is the expected value of roll angle,/>As a measure of roll angle,/>Is a proportional control parameter in a roll angle proportional integral controller,/>Is an integral control parameter in a roll angle proportional integral controller,/>For proportional control parameters in roll angle speed proportional integral controller,/>For integral control parameters in roll angle speed proportional integral controller,/>Is the deflection angle of aileron,/>The trim angle for the aileron;
yaw angle control loop: the yaw angle and yaw rate controller in the loop is designed to:
wherein, Is the expected value of yaw rate,/>For yaw rate measurement,/>Is the expected value of yaw angle,/>Is a measurement of yaw angle,/>For proportional control parameters in yaw proportional integral controller,/>For integral control parameters in yaw proportional integral controller,/>For proportional control parameters in yaw rate proportional integral controller,/>For integral control parameters in yaw rate proportional integral controller,/>Is the deflection angle of the rudder,/>The trim angle for the rudder.
5. A method of high reliability and safety flight control for enhancing the stability reserve of a system as claimed in claim 4, wherein: in step 2, the aircraft attitude angle control system for enhancing the stable reserve comprises:
Pitch angle control loop: the pitch angle, pitch angle speed and pitch angle acceleration controller in the loop are designed as follows:
wherein, Is the expected value of pitch angle acceleration signal,/>Is the estimated value of pitch angle acceleration signal,/>For proportional control parameters in a pitch angle acceleration proportional integral controller,/>Integrating control parameters of a pitch angle acceleration proportional integral controller;
Roll angle control circuit: the roll angle, the roll angle speed and the roll angle acceleration controller in the loop are designed as follows:
wherein, Is the expected value of the roll angle acceleration signal,/>Is the estimated value of the roll angle acceleration signal,/>For proportional control parameters in roll angle acceleration proportional integral controller,/>Integrating control parameters in a roll angle acceleration proportional integral controller;
yaw angle control loop: yaw angle, yaw rate and yaw acceleration controllers in the circuit are designed as follows:
wherein, Is the expected value of yaw acceleration signal,/>For the estimated value of yaw acceleration signal,/>For proportional control parameters in yaw acceleration proportional integral controller,/>And integrating control parameters of a proportional-integral controller for yaw angular acceleration.
6. A highly reliable and safe flight control method for enhancing system stability reserve according to claim 4 or 5, wherein: in step 2, sensor noise is introduced into attitude angle measurement values and attitude angular speed measurement values in the aircraft attitude angle control system for enhancing the stable reserve.
7. A method of high reliability and safety flight control for enhancing the stability reserve of a system as claimed in claim 6, wherein: the attitude angle sensor noise adopts a Gaussian white noise model, and the time delay of the angle sensor of 40e-3 seconds is introduced after the noise is added; the noise of the attitude angular velocity sensor adopts a Gaussian white noise model, the speed limit of 600deg/s is carried out on the attitude angular velocity, and the time delay of the angular velocity sensor of 15e-3 seconds is introduced after the noise is added.
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