CN112000119B - Aircraft lateral overload tracking control method taking attitude stabilization as core - Google Patents

Abstract

The application discloses an aircraft lateral overload tracking control method taking stable posture as a core, which is characterized in that an angular accelerometer is adopted to measure the lateral overload of an aircraft, overload errors are obtained by subtracting an overload instruction from the lateral overload of the aircraft, nonlinear integration is carried out on the overload errors, and an overload error integral and an overload error signal are superposed to form a comprehensive signal. And then performing nonlinear integral operation and superposing proportional integral signals of the integrated signals to form a desired yaw angle signal of the aircraft, and driving an attitude stabilizing loop of the aircraft. And the attitude stabilizing loop adopts proportional, integral and nonlinear differential compound control of yaw angle errors. The finally obtained control law signal drives an aircraft yaw channel rudder system to control the aircraft course plane to turn and fly, and the aim of tracking the expected overload instruction signal by the lateral overload of the aircraft yaw channel is fulfilled. The application can give consideration to the advantages of large stability margin of the gesture and good maneuverability of overload control.

Description

An aircraft lateral overload tracking control method with attitude stability as the core

Technical field

The invention belongs to the field of aircraft control, and particularly relates to the design of attitude control loops and overload control loops for stable flight of the aircraft. It can also be applied to the terminal guidance loop of the aircraft using proportional guidance.

Background technique

Aircraft stability control is the core technology of aircraft flight. Currently, the more popular methods include attitude stability control system and overload stability control system. Moreover, most aircraft mainly use attitude stabilization control methods, especially manned aircraft, mainly because the attitude stabilization control method has a good stability margin, and after many years of extensive application, more design and use experience has been accumulated. Hence also better reliability. However, overload stability control is used in unmanned aerial vehicles that pursue high maneuverability. The advantage is that aircraft using an overload control system have better turning speed and maneuverability. While traditional overload control generally no longer measures the attitude angle of the aircraft, the present invention proposes a design method for an aircraft overload control loop with attitude control as the core. Its main advantage is that it can retain the good stability of attitude control and has overload capability. The advantages of rapid control and good maneuverability. It is especially suitable for use in certain unmanned aerial vehicle controls that require switching from attitude control to overload control for proportional guidance at the end of flight. Therefore, the present invention has high engineering application value and can be widely used in the field of aircraft control.

It should be noted that the information disclosed in the above background section is only used to enhance the understanding of the background of the present invention, and therefore may include information that does not constitute prior art known to those of ordinary skill in the art.

Contents of the invention

The purpose of the present invention is to provide an aircraft lateral overload tracking control method with attitude stability as the core, and thereby overcome, at least to a certain extent, the problem of overload control with sufficient maneuverability and insufficient stability margin due to limitations and defects in related technologies. question.

According to one aspect of the present invention, an aircraft lateral overload tracking control method with attitude stability as the core is provided, including the following steps:

Step S10: Install an angle gyroscope and a rate gyroscope on the aircraft to measure the yaw angle and yaw angle rate of the aircraft, and use a linear accelerometer to measure the lateral overload of the aircraft;

Step S20: Compare the overload signal with the overload command signal to obtain an overload error signal, perform linear and nonlinear integration respectively, and superimpose it with the overload error signal to form an outer loop comprehensive signal;

Step S30: Perform linear integral and nonlinear integral operations on the outer loop comprehensive signal to obtain the integral and nonlinear integral signal of the comprehensive signal, and then perform signal synthesis to obtain the desired yaw angle signal;

Step S40: Compare the desired yaw angle signal with the yaw angle measurement signal to obtain a yaw angle error signal, then perform anti-saturation nonlinear transformation on the yaw angle rate signal to obtain a comprehensive yaw angle rate signal, and finally perform The final yaw channel control signal is obtained comprehensively;

Step S50: According to the design of the yaw channel control signal, conduct debugging parameters and command tracking tests, complete the attitude stabilization loop design of the yaw channel, realize the tracking of the desired yaw angle, and at the same time complete the tracking of the desired lateral overload signal. .

In an exemplary embodiment of the present invention, the overload error signal is obtained by comparing the overload signal and the overload command signal, linear and nonlinear integration are performed respectively, and superimposed with the overload error signal to form an outer loop comprehensive signal including;

s _e1 =∫e _n dt;

u ₁ ＝k ₁ e _n +k ₂ s _e1 +k ₃ s _e2 ;

in represents the overload command test signal, n _z represents the lateral overload measurement value of the aircraft, and _en represents the overload error signal. s _e1 is the outer overload error integral signal, s _e2 is the overload error nonlinear integral term, ε ₁ is the positive control parameter, k ₁ , k ₂ and k ₃ are control parameters. For detailed selection, see the case implementation below. u ₁ is the comprehensive signal of the outer loop.

In an exemplary embodiment of the present invention, linear integral and nonlinear integral operations are performed on the outer loop comprehensive signal to obtain the integral and nonlinear integral signal of the comprehensive signal, and then the signal is synthesized to obtain the desired yaw angle signal. include:

s _u1 =∫u ₁ dt;

ψ ^d ＝k ₄ u ₁ +k ₅ s _u1 +k ₆ s _u2 ;

Among them, u ₁ is the comprehensive signal of the outer loop, s _u1 is the integrated signal of the integrated signal, s _u2 is the nonlinear integral signal of the integrated signal, ε ₂ is the positive control parameter, k ₄ , k ₅ and k ₆ are the control parameters. The details Select the case implementation shown below, and ψ ^d is the expected yaw angle signal.

In an exemplary embodiment of the present invention, the desired yaw angle signal is compared with the yaw angle measurement signal to obtain a yaw angle error signal, and then the yaw angle rate signal is subjected to anti-saturation nonlinear transformation to obtain the yaw angle rate signal. The angle rate synthesis signal is finally synthesized to obtain the final yaw channel control signal including:

e _ψ = ψ ^d -ψ;

s _ψ =∫e _ψ dt;

u＝k ₈ e _ψ +k ₉ s _ψ +k ₁₀ D _ψ ;

Among them, ψ ^d is the desired yaw angle signal, ψ is the yaw angle measurement signal of the aircraft, e _ψ is the yaw angle error signal, s _ψ is the yaw angle error integrated signal, ω _y is the yaw angle rate signal, D _ψ is the comprehensive signal of yaw angular rate, k ₇ , k ₈ , k ₉ , k ₁₀ and ε ₃ are control parameters. Please refer to the case implementation later for detailed selection. u is the final yaw channel control signal.

beneficial effects

The present invention provides a new overload control method. Unlike traditional overload control which only measures overload without paying attention to attitude stability, the overload stability of the present invention is based on attitude stability, so the final signal is used to drive the attitude. Stable loop. Such a method helps to use the traditional aircraft attitude stabilization loop to realize the proportional guidance of the terminal stage, which not only maintains the advantages of the strong stability of the attitude stabilization loop, but also can be combined with the proportional guidance of the terminal aircraft, using Overload control realizes proportional guidance. At the same time, the overload error internal and external loop design method was proposed, especially the nonlinear integration method, which solved the signal matching problem between the overload and attitude loops, and achieved accurate tracking of the overload signal without static error. Therefore, the method proposed in this article has high engineering application value.

It should be understood that the above general description and the following detailed description are exemplary and explanatory only, and do not limit the present invention.

Description of the drawings

The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments consistent with the invention and together with the description serve to explain the principles of the invention. Obviously, the drawings in the following description are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without exerting creative efforts.

Figure 1 is a flow chart of an aircraft lateral overload tracking control method with attitude stability as the core provided by the present invention;

Figure 2 is the aircraft yaw angle curve (unit: degree) of the method provided by the case of the present invention;

Figure 3 is the aircraft yaw angle rate curve (unit: degree/second) of the method provided by the case of the present invention;

Figure 4 is the aircraft lateral overload curve (unit: g) of the method provided by the case of the present invention;

Figure 5 is the aircraft lateral overload error curve (unit: g) of the method provided by the case of the present invention;

Figure 6 is the nonlinear integral of the lateral overload error (unitless) of the method provided by the case of the present invention;

Figure 7 is the integrated signal u1 curve (unitless) of the method provided by the case of the present invention;

Figure 8 is the nonlinear integral curve (unitless) of the integrated signal u1 of the method provided by the case of the present invention;

Figure 9 is the expected yaw angle curve (unit: degree) of the aircraft according to the method provided by the case of the present invention;

Figure 10 is the final control quantity curve of the aircraft (unitless) of the method provided by the case of the present invention;

Figure 11 is the aircraft sideslip angle curve (unit: degree) of the method provided by the case of the present invention;

Figure 12 is the aircraft yaw and rudder angle curve (unit: degree) of the method provided by the case of the present invention.

Detailed ways

Example embodiments will now be described more fully with reference to the accompanying drawings. Example embodiments may, however, be embodied in various forms and should not be construed as limited to the examples set forth herein; rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the concepts of the example embodiments. To those skilled in the art. The described features, structures or characteristics may be combined in any suitable manner in one or more embodiments. In the following description, numerous specific details are provided to provide a thorough understanding of the embodiments of the invention. However, those skilled in the art will appreciate that the technical solutions of the present invention may be practiced without one or more of the specific details described, or other methods, components, devices, steps, etc. may be adopted. In other instances, well-known technical solutions have not been shown or described in detail to avoid obscuring aspects of the present invention.

The present invention provides an aircraft lateral overload tracking control method with attitude stability as the core. It uses an angular accelerometer to measure the lateral overload of the aircraft. The overload error is obtained by subtracting the aircraft lateral overload and the overload command, and then the overload error is obtained. Nonlinear integration is performed, and the overload error integral and the overload error signal are superimposed to form a comprehensive signal. On the basis of the comprehensive signal, a nonlinear integral operation is performed here and the proportional integral signal of the comprehensive signal is superimposed to form the desired yaw angle signal of the aircraft and drive the aircraft attitude stabilization loop. The attitude stabilization loop uses proportional, integral, and nonlinear differential composite control of the yaw angle error. The finally obtained control law signal drives the aircraft's yaw channel rudder system, controls the aircraft's heading and plane turning flight, and achieves the purpose of the aircraft's yaw channel lateral overload tracking of the desired overload command signal.

Below, the aircraft lateral overload tracking control method with attitude stability as the core according to the present invention will be further explained and described in conjunction with the accompanying drawings. Referring to Figure 1, this aircraft lateral overload tracking control method with attitude stability as the core includes the following steps:

Step S10: Install an angle gyroscope and a rate gyroscope on the aircraft to measure the yaw angle and yaw angle rate of the aircraft, and use a linear accelerometer to measure the lateral overload of the aircraft;

An angle gyroscope is used to measure the yaw angle of the aircraft, which is recorded as ψ. A rate gyroscope is used to measure the yaw rate of the aircraft, denoted as ω _y . An accelerometer is used to measure the lateral overload of the aircraft, which is recorded as n _z . The above-mentioned angle gyroscope, angular rate gyroscope and accelerometer are all installed on the axis of the aircraft body.

Step S20: Compare the overload signal with the overload command signal to obtain an overload error signal, perform linear and nonlinear integration respectively, and superimpose it with the overload error signal to form an outer loop comprehensive signal;

Specifically, first set the overload command test signal to a constant value, denoted as See the case implementation below for its specific selection value. Then, the aircraft's lateral overload measurement value n _z and the above-mentioned overload command test signal/> After comparison, the obtained signal is called the overload error signal, denoted as _en . The comparison method is subtraction as follows:/>

Secondly, perform an integral operation on the above error signal to obtain the outer layer overload error integral signal, which is denoted as s _e1 . The integral operation is performed according to the following formula:

s _e1 =∫e _n dt;

Then design the overload error nonlinear integral term, denoted as s _e2 , which is calculated as follows:

Among them, ε ₁ is a positive control parameter, and its detailed selection is shown in the case implementation below.

Finally, the above three types of signals are proportionally superimposed to obtain the comprehensive signal of the outer loop, denoted as u ₁ , which is defined as:

u ₁ ＝k ₁ e _n +k ₂ s _e1 +k ₃ s _e2 ;

Among them, k ₁ , k ₂ and k ₃ are control parameters. For their detailed selection, see the case implementation below.

Step S30: Perform linear integral and nonlinear integral operations on the outer loop comprehensive signal to obtain the integral and nonlinear integral signal of the comprehensive signal, and then perform signal synthesis to obtain the desired yaw angle signal;

In this step, proportional and nonlinear integral operations are mainly performed on the outer loop comprehensive signal u ₁ to obtain the inner loop signal. Specifically, the comprehensive signal is first integrated, which is called the integrated signal, denoted as s _u1 , and its integral operation is performed according to the following formula:

s _u1 =∫u ₁ dt;

Secondly, perform nonlinear integration on the comprehensive signal u ₁ to obtain the nonlinear integral signal of the comprehensive signal, which is denoted as s _u2 . The integral operation is performed according to the following formula:

Among them, ε ₂ is a positive control parameter, and its detailed selection is shown in the case implementation below.

Finally, the above signals are synthesized again, and the obtained inner loop signal is used as the driving signal of the attitude stabilization loop, which is called the desired yaw angle signal, denoted as ψ ^d , and its calculation is carried out according to the following formula

ψ ^d ＝k ₄ u ₁ +k ₅ s _u1 +k ₆ s _u2 ;

Among them, k ₄ , k ₅ and k ₆ are control parameters. For their detailed selection, see the case implementation below.

Step S40: Compare the desired yaw angle signal with the yaw angle measurement signal to obtain a yaw angle error signal, then perform anti-saturation nonlinear transformation on the yaw angle rate signal to obtain a comprehensive yaw angle rate signal, and finally perform The final yaw channel control signal is obtained comprehensively;

Specifically, the above-mentioned expected yaw angle signal ψ ^d is first used to compare with the yaw angle measurement signal ψ of the aircraft. The obtained signal is called the yaw angle error signal, denoted as e _ψ . The comparison method is as follows: e _ψ = ψ ^d -ψ.

Secondly, the above yaw angle error is integrated to obtain the yaw angle error integrated signal, which is recorded as s _ψ . The integral is calculated according to the following formula:

s _ψ =∫e _ψ dt;

Thirdly, the yaw angle rate signal measured by the gyroscope is proportioned and anti-saturation nonlinear changes are made, which is called the yaw angle rate comprehensive signal, denoted as D _ψ , and its calculation is as follows:

Among them, k ₇ and ε ₃ are control parameters. For detailed selection, please refer to the case implementation below.

Finally, the above-mentioned yaw angle error signal e _ψ , yaw angle error integrated signal s _ψ and yaw angle rate comprehensive signal D _ψ are linearly synthesized to obtain the final yaw channel control signal, denoted as u, which is calculated as follows:

u＝k ₈ e _ψ +k ₉ s _ψ +k ₁₀ D _ψ ;

Among them, k ₈ , k ₉ , and k ₁₀ are control parameters. For their detailed selection, see the case implementation below.

Step S50: According to the design of the yaw channel control signal, conduct debugging parameters and command tracking tests, complete the attitude stabilization loop design of the yaw channel, realize the tracking of the desired yaw angle, and at the same time complete the tracking of the desired lateral overload signal. .

Follow the above method to build the final control signal u and output it to the aircraft yaw rudder system to control the lateral movement of the aircraft to achieve the aircraft's lateral overload n _z to track the desired overload command test signal By comparing n _z with/> curve, if the tracking situation meets the requirements, the design of the lateral overload control loop is completed; if not, the control parameters are adjusted and redesigned until the appropriate control parameters are selected and the design is completed.

Case implementation and computer simulation results analysis

In order to verify the correctness and effectiveness of the method provided by the present invention, the following case simulation is provided for simulation.

In step one, the main purpose is to measure the status of the aircraft. Since this case mainly illustrates the effectiveness of the overload control method provided by the present invention, in this case design, the tracking of the overload command starts 5 seconds after the aircraft is launched. The main purpose is to avoid the acceleration process of the aircraft during the launch phase and the interference of unstable speed on the system, so as to avoid difficulties in parameter selection in the overload control loop caused by speed changes. We use an angle gyroscope to measure the yaw angle ψ of the aircraft. In this case, the yaw angle of the aircraft flying for 15 seconds is shown in Figure 2 below. A rate gyroscope is used to measure the yaw angular rate ω _y of the aircraft. The yaw angular rate curve of the entire 15-second flight is shown in Figure 3 below. An accelerometer is used to measure the lateral overload of the aircraft. The lateral overload curve of the entire 15-second flight is shown in Figure 4.

In step two, the main purpose is to design the outer loop of overload error and integration. We set the overload command test signal n _y ^* = 1 and compare it with the lateral overload measurement value n _z and the obtained overload error signal e _n The curve is shown in Figure 5. Select ε ₁ =0.5, and the curve of the overload error nonlinear integral term s _e2 is shown in Figure 6. Select k ₁ =-2, k ₂ =-4 and k ₃ =-5, and the curve of the integrated signal u ₁ is shown in Figure 7.

In step three, the main purpose is to design the inner loop of overload error and integration. We select ε ₂ =2. The variation curve of the nonlinear integral signal s _u2 of the integrated signal u ₁ is shown in Figure 8. Selecting k ₄ =5, k ₅ =1.7 and k ₆ =5, the final expected yaw angle signal ψ ^d variation curve is shown in Figure 9.

In step four, the attitude stabilization loop design is mainly carried out. k ₇ =3 and ε ₃ =20 are selected, k ₄ =0.35, k ₅ =0.25 and k ₆ =0.15. The final control quantity u is obtained as shown in Figure 10 .

In step five, the main focus is on debugging parameters and instruction tracking testing. According to the above parameters, it can be seen that the final aircraft lateral overload can stably track the expected overload command test signal. At the same time, adjust the size and amplitude of the instructions and find that the system can track stably. If the flight environment, such as altitude and speed, changes, some parameters need to be adjusted to achieve a satisfactory dynamic process, so that the parameters can be selected and the design completed. The final rudder angle curve of the aircraft is shown in Figure 11, and the sideslip angle of the aircraft is shown in Figure 12. It can be seen that the sideslip angle of the aircraft does not exceed the available range by 12 degrees, and the rudder deflection angle does not exceed the available range by 25 degrees. Therefore, it shows that the lateral overload control method provided by the present invention is reasonable and effective.

Other embodiments of the invention will be readily apparent to those skilled in the art from consideration of the specification and practice of this type of invention. This application is intended to cover any variations, uses, or adaptations of the invention that follow the general principles of the invention and include common knowledge or customary technical means in the technical field not specified in the invention. . It is intended that the specification and examples be considered as exemplary only, with a true scope and spirit of the invention being indicated by the following claims.

Claims (1)

1. An aircraft lateral overload tracking control method taking attitude stabilization as a core is characterized by comprising the following steps:

step S10: installing an angle gyroscope and a rate gyroscope on an aircraft, measuring the yaw angle and the yaw rate of the aircraft, and simultaneously measuring the lateral overload of the aircraft by adopting a linear accelerometer;

step S20: and comparing the lateral overload signal with an overload instruction signal to obtain an overload error signal, respectively performing linear and nonlinear integration, and overlapping the overload error signal to form an outer loop integrated signal as follows:

S _e1 ＝∫e _n dt；

u ₁ ＝k ₁ e _n +k ₃ s _e1 +k ₃ s _e2 ；

wherein the method comprises the steps ofFor overload command test signal, n _z E, for the lateral overload measurement of the aircraft _n Is an overload error signal; s is(s) _e1 Integrating the signal for the outer layer overload error s _e2 Epsilon is the nonlinear integral term of overload error ₁ Is a positive control parameter, k ₁ 、k ₂ And k is equal to ₃ Is a control parameter; u (u) ₁ Is an outer loop integrated signal;

step S30: and performing linear integration and nonlinear integration operation on the outer loop integrated signal to obtain integrated and nonlinear integrated signals of the integrated signal, and performing signal synthesis to obtain a desired yaw angle signal as follows:

S _u1 ＝∫u ₁ dt；

ψ ^d ＝k ₄ u ₁ +k ₅ s _u1 +k ₆ s _u2 ；

wherein u is ₁ Is the outer loop integrated signal s _u1 Integrate the signals for the combined signal s _u2 Epsilon is the non-linear integral of the integrated signal ₂ Is a positive control parameter, k ₄ 、k ₅ And k is equal to ₆ To control parameters, ψ ^d A desired yaw angle signal;

step S40: comparing the expected yaw angle signal with the yaw angle measurement signal to obtain a yaw angle error signal, performing anti-saturation nonlinear transformation on the yaw angle rate signal to obtain a yaw angle rate integrated signal, and finally integrating to obtain a final yaw channel control signal as follows:

e _ψ ＝ψ ^d -ψ；

S _ψ ＝∫e _ψ dt；

u＝k ₈ e _ψ +k ₉ s _ψ +k ₁₀ D _ψ ；

wherein, psi is ^d For the desired yaw angle signal, ψ is the yaw angle measurement signal of the aircraft, e _ψ Is a yaw angle error signal s _ψ To integrate the yaw angle error signal omega _y For yaw rate signal, D _ψ For yaw rate integrated signal, k ₇ 、k ₈ 、k ₉ 、k ₁₀ And epsilon ₃ U is a final yaw channel control signal for the control parameter;

step S50: and according to the design of the yaw channel control signal, debugging parameters and instruction tracking tests are carried out, the design of a stable attitude loop of the yaw channel is completed, the tracking of an expected yaw angle is realized, and the tracking of an expected lateral overload signal is completed.