Rolling aircraft control method based on two-loop roll attitude pilot

Abstract

The application discloses a roll aircraft control method based on a two-loop roll attitude pilot, in which a roll angle error and a roll angle rate are obtained in real time, a roll rudder deflection control instruction corresponding to a unit roll angle error and a roll attitude pilot damping coefficient are further weighted, so that a rudder deflection control instruction is obtained and is transmitted to a rudder, the rudder works based on the rudder deflection control instruction, and the roll aircraft is controlled to fly to a target.

Description

Technical Field

[0001] This invention relates to a control method for a roll aircraft, specifically a control method for a roll aircraft based on a two-loop roll attitude autopilot. Background Technology

[0002] The design purpose of autopilots for roll-type aircraft is to robustly track input commands by generating stable responses through measurement. The requirements of a new generation of short-range aircraft for high angle of attack and high agility have driven research into guidance and autopilot design.

[0003] The difference between roll attitude control systems and pitch / yaw attitude control systems is that the design of roll attitude control systems must consider the inherent disturbance torque in the roll channel. This disturbance torque primarily arises when the plane of attack at full angle of attack is not on the aircraft's plane of symmetry. Due to the asymmetry of the maneuvering surfaces, a disturbance roll torque is generated. The system design must ensure that the steady-state roll attitude error does not exceed the allowable value under the maximum possible disturbance roll torque. When designing a two-loop roll attitude control system, the design parameters in the transfer function must be determined to obtain the roll rudder deflection control command and the roll attitude control system damping coefficient corresponding to a unit roll angle error. However, current common solutions simply set this design parameter to a fixed value or a time-varying parameter, leading to a deviation from reality and affecting the subsequent hit accuracy of the aircraft.

[0004] Based on this, the inventors have conducted in-depth research on the control of rolling aircraft, especially the specific value scheme of the design parameters, in order to design a rolling aircraft control method based on a two-loop rolling attitude autopilot that can solve the above problems. Summary of the Invention

[0005] To overcome the above problems, the inventors conducted in-depth research and designed a control method for a rolling aircraft based on a two-loop rolling attitude autopilot. In this method, the roll angle error and roll angle rate are obtained in real time. Furthermore, the roll rudder deflection control command corresponding to the unit roll angle error and the damping coefficient of the rolling attitude autopilot are weighted and processed to obtain the rudder deflection control command, which is then transmitted to the servo motor. The servo motor operates based on the rudder deflection control command to control the rolling aircraft to fly toward the target, thus completing the invention.

[0006] Specifically, the purpose of this invention is to provide a control method for a rolling aircraft based on a two-loop rolling attitude autopilot. In this method, the roll angle error and roll angle rate are obtained in real time, and then the rudder deflection control command is obtained through weighted processing and transmitted to the servo motor. The servo motor operates based on the rudder deflection control command to control the rolling aircraft to fly toward the target.

[0007] The rudder deflection control command is obtained through the following formula (I):

[0008]

[0009] Where Δγ represents the roll angle error;

[0010] Indicates the roll rate;

[0011] K A This represents the roll rudder deflection control command corresponding to a unit roll angle error.

[0012] K g This indicates the damping coefficient of the roll attitude control system.

[0013] Wherein, the roll rate We obtain it through the following formula (ii):

[0014]

[0015] Where s represents the sign of the Laplace transform;

[0016] k r and T r Each represents a dynamic coefficient independently;

[0017]

[0018]

[0019] This represents the derivative of roll damping;

[0020] This represents the derivative of the rolling torque generated by the rudder;

[0021] J x This represents the moment of inertia of an aircraft about its longitudinal axis.

[0022] Wherein, the k A Obtained through the following formula (iii):

[0023]

[0024] Where, ω CR This indicates the open-loop crossing frequency index of the driving instrument;

[0025] c δ Indicates the dynamic coefficient;

[0026]

[0027] Wherein, K g Obtained through the following formula (iv):

[0028]

[0029] Where ζ represents the damping coefficient;

[0030] ω CR This indicates the open-loop crossing frequency index of the driving instrument;

[0031] c δ and c ω Each represents a dynamic coefficient independently;

[0032]

[0033]

[0034] The beneficial effects of this invention include:

[0035] The rolling aircraft control method based on a two-loop rolling attitude autopilot provided by the present invention constructs a roll rudder deflection control command and a roll attitude autopilot damping coefficient corresponding to a unit roll angle error based on the autopilot's open-loop crossover frequency index and damping coefficient. This allows for more accurate weighted processing of the roll angle error and roll angle rate, providing more accurate rudder deflection control commands for subsequent servo control and improving the aircraft's hit accuracy. Attached Figure Description

[0036] Figure 1 A schematic diagram of the step response obtained in the embodiment is shown;

[0037] Figure 2 The root locus diagram obtained in the embodiment is shown;

[0038] Figure 3 A schematic diagram of the line-of-sight angle estimation error obtained in the embodiment is shown;

[0039] Figure 4 A schematic diagram of the step response obtained in the comparative example is shown;

[0040] Figure 5 The root locus obtained in the comparative example is shown;

[0041] Figure 6 A schematic diagram showing the line-of-sight angle estimation error obtained in the comparative example is shown. Detailed Implementation

[0042] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. Through these descriptions, the features and advantages of the present invention will become clearer and more apparent.

[0043] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments. Although various aspects of embodiments are shown in the accompanying drawings, the drawings are not necessarily drawn to scale unless specifically indicated otherwise.

[0044] This application provides a control method for a rolling aircraft based on a two-loop rolling attitude autopilot. In this method, the roll angle error and roll angle rate are obtained in real time, and then the rudder deflection control command is obtained through weighted processing and transmitted to the servo motor. The servo motor operates based on the rudder deflection control command to control the rolling aircraft to fly toward the target.

[0045] In a preferred embodiment, the rudder deflection control command is obtained by the following formula (a):

[0046]

[0047] Where Δγ represents the roll angle error; in this application, the roll angle γ of the aircraft is obtained in real time through the inertial navigation unit on the aircraft. c Then, the roll angle γ will be obtained. c The roll angle error is obtained by subtracting the roll angle γ output by the driver's instrument.

[0048] Indicates the roll rate;

[0049] K A This represents the forward gain of the roll attitude control system, which is the roll rudder deflection control command corresponding to a unit roll angle error, used to control the attitude angle of the aircraft.

[0050] K g This represents the damping coefficient of the roll attitude control system, and the roll rudder deflection control command corresponding to a unit roll angular rate. It is used to improve the roll direction damping characteristics of the aircraft and increase roll damping.

[0051] In this application, the parameter K is solved. A and K g This helps improve the damping characteristics of the aircraft's attitude angle and roll direction.

[0052] Preferably, the roll rate We obtain it through the following formula (ii):

[0053]

[0054] Where s represents the sign of the Laplace transform;

[0055] k r and T r Each represents a dynamic coefficient independently;

[0056]

[0057]

[0058] This represents the derivative of roll damping, whose value is determined based on the aircraft's hardware parameters before launch.

[0059] The derivative of the rolling torque generated by the rudder is determined based on the aircraft's hardware parameters before launch.

[0060] J x It represents the moment of inertia of the aircraft about its longitudinal axis. Its value is determined based on the aircraft's hardware parameters before launch.

[0061] Preferably, the k A We obtain it through the following formula (iii):

[0062]

[0063] Where, ω CR This refers to the open-loop cross-frequency index of the autopilot, whose value is determined based on the aircraft's hardware parameters before launch.

[0064] c δ Indicates the dynamic coefficient;

[0065]

[0066] Preferably, the kg is obtained by the following formula (iv):

[0067]

[0068] Where ζ represents the damping coefficient, the value of which is determined based on the hardware parameters of the aircraft before launch;

[0069] ω CR This indicates the open-loop crossing frequency index of the driving instrument;

[0070] c δ and c ω Each represents a dynamic coefficient independently;

[0071]

[0072]

[0073] Example

[0074] The following values ​​are set for the power coefficient and related airframe parameters of a certain type of roll aircraft:

[0075] ω CR = 40 rad / s

[0076] ζ = 0.7

[0077]

[0078]

[0079] J x =0.96 kg·m 2

[0080] To control the aircraft to intercept a maneuvering target 10km away, the guidance rate used is the guidance rate provided in Chinese Patent 2019103567261, "Low-Cost Integrated Navigation System and Method Applicable to High-Speed ​​Rolling Aircraft." The specific control process is as follows:

[0081] The roll angle error and roll rate are obtained in real time. The weighted processing is then used to obtain the rudder deflection control command and transmit it to the servo motor. The servo motor then operates the rudder based on the rudder deflection control command to control the rolling aircraft to fly toward the target.

[0082] The rudder deflection control command is obtained through the following formula (i):

[0083]

[0084] Where Δγ represents the roll angle error, which is calculated by adjusting the roll angle γ. c It is obtained by subtracting the roll angle γ output by the driver's instrument.

[0085] Indicates the roll rate;

[0086] The roll rate We obtain it through the following formula (ii):

[0087]

[0088] Where s represents the sign of the Laplace transform;

[0089] k r and T r Each represents a dynamic coefficient independently;

[0090]

[0091]

[0092] The K A Obtained through the following formula (iii):

[0093]

[0094] The K g Obtained through the following formula (iv):

[0095]

[0096] pass Get c δ =7276.69287, c ω =31.95242;

[0097] Finally, K was obtained. A =0.2199, K g =0.0033.

[0098] The obtained K A and K g Substituting the servo transfer function into the simulation (which is set to -1 in this simulation), and performing the simulation using Simulink, the simulation results are as follows. Figure 1 , Figure 2 and Figure 3 As shown in the image.

[0099] pass Figure 1 It can be seen that the step response time is short, which is related to the small damping coefficient of the aircraft, and the curve converges quickly to 1.

[0100] pass Figure 2 It can be seen that all the zeros and poles are on the left side, meaning the system is stable.

[0101] Bode plots drawn based on open-loop transfer functions, such as Figure 3 As shown in the diagram, the servo transfer function is set to -1, representing an ideal state. However, in reality, the servo is a second-order circuit. Figure 3 The obtained ω CR = 38.2 rad / s, compared to the actual ω CR =40rad / s is close, and the error is within a reasonable range.

[0102] Comparative Example

[0103] Using the exact same aircraft as in Example 1, targeting the exact same maneuvering target, and employing the exact same guidance rate, the control process is as follows:

[0104] The roll angle error and roll rate are obtained in real time. The weighted processing is then used to obtain the rudder deflection control command and transmit it to the servo motor. The servo motor then operates the rudder based on the rudder deflection control command to control the rolling aircraft to fly toward the target.

[0105] The rudder deflection control command is obtained through the following formula (i):

[0106]

[0107] Where Δγ represents the roll angle error, obtained by obtaining the roll angle γ c Obtained by subtracting the roll angle γ output by the driver's instrument;

[0108] Indicates the roll rate;

[0109] The roll rate We obtain it through the following formula (ii):

[0110]

[0111] Where s represents the symbol of the Laplace transform.

[0112] k r and T r Each represents a dynamic coefficient independently;

[0113]

[0114]

[0115] The K A The value is 1.2;

[0116] The K g The value is 0.004;

[0117] The final simulation results are as follows Figure 4 , Figure 5 and Figure 6 As shown;

[0118] pass Figure 4 It can be seen that the curve oscillates severely and gradually diverges, failing to converge to 1.

[0119] pass Figure 5 It can be seen that not all zeros and poles are on the left side, indicating that the system is unstable.

[0120] Bode plots drawn based on open-loop transfer functions, such as Figure 6 As shown, the open-loop crossover frequency ω of the autopilot CR The corresponding phase margin is less than 0, indicating that the system is unstable.

[0121] The results from the embodiments and comparative examples show that the roll aircraft control method based on a two-loop roll attitude autopilot in the embodiments can ensure stable flight of the aircraft and more accurate target hits.

[0122] The present invention has been described above with reference to preferred embodiments; however, these embodiments are merely exemplary and illustrative. Various substitutions and modifications can be made to the present invention based on these embodiments, all of which fall within the scope of protection of the present invention.

Claims

1. A control method for a roll aircraft based on a two-loop roll attitude autopilot, characterized in that, In this method, the roll angle error and roll rate are obtained in real time, and then the rudder deflection control command is obtained through weighted processing and transmitted to the servo motor. The servo motor operates based on the rudder deflection control command to control the rolling aircraft to fly toward the target. The rudder deflection control command is obtained through the following formula (i): Where Δγ represents the roll angle error; Indicates the roll rate; K A This represents the roll rudder deflection control command corresponding to a unit roll angle error. K g This indicates the damping coefficient of the roll attitude control system; The K A Obtained through the following formula (iii): Where, ω CR This indicates the open-loop crossing frequency index of the driving instrument; c δ Indicates the dynamic coefficient; The K g Obtained through the following formula (iv): Where ζ represents the damping coefficient; ω CR This indicates the open-loop crossing frequency index of the driving instrument; c δ and c ω Each represents a dynamic coefficient independently; This represents the derivative of roll damping; This represents the derivative of the rolling torque generated by the rudder; J x This represents the moment of inertia of an aircraft about its longitudinal axis.

2. The roll aircraft control method based on a two-loop roll attitude autopilot according to claim 1, characterized in that, The roll rate We obtain it through the following formula (ii): Where s represents the sign of the Laplace transform; k r and T r Each represents a dynamic coefficient independently;

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