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

Abstract

The invention discloses a rolling aircraft control method based on a two-loop rolling attitude pilot, and the method comprises the steps: obtaining a rolling angle error and a rolling angle rate in real time, and further carrying out the weighting processing through a rolling rudder deviation control instruction corresponding to a unit rolling angle error and a damping coefficient of the rolling attitude pilot, therefore, a rudder deflection control instruction is obtained and transmitted to the steering engine, and the steering engine performs steering work based on the rudder deflection control instruction to control the rolling aircraft to fly to a target.

Description

Technical Field

[0001] The invention relates to a rolling aircraft control method, and in particular to a rolling aircraft control method based on a two-loop rolling attitude autopilot. Background Art

[0002] The design of autopilot for rolling aircraft aims to use the measurement to produce a stable response to robustly track the input command. The requirements of a new generation of short-range aircraft for high angle of attack and high agility have promoted the research on guidance and autopilot design.

[0003] The difference between the roll attitude autopilot and the pitch / yaw attitude autopilot is that its design must take into account the interference torque that is fixed in the roll channel. Its main source is that when the full angle of attack action surface is not on the symmetric surface of the aircraft, the interference roll torque will be generated due to the left-right asymmetry of the maneuvering surface. The autopilot should be designed to ensure that the roll attitude steady-state angle error does not exceed the allowable value under the maximum possible interference roll torque. When designing a two-loop roll attitude autopilot, it is necessary to determine the design parameters in the transfer function, and then obtain the roll rudder deflection control command and roll attitude autopilot damping coefficient corresponding to the unit roll angle error. However, in the current common schemes, the design parameters are simply set to fixed values ​​or time-varying parameters, resulting in a certain deviation from the actual situation, affecting the subsequent aircraft's hit accuracy.

[0004] Based on this, the inventors have conducted in-depth research on the control of the rolling vehicle, especially the specific value schemes of the design parameters, in the hope of designing a rolling vehicle control method based on a two-loop roll attitude autopilot that can solve the above-mentioned problems. Summary of the invention

[0005] In order to overcome the above-mentioned problems, the inventors have conducted intensive research and designed a rolling aircraft control method based on a two-loop roll attitude autopilot. In this method, the roll angle error and the roll angle rate are obtained in real time, and further weighted processing is performed through the roll rudder control instruction corresponding to the unit roll angle error and the roll attitude autopilot damping coefficient, so as to obtain the rudder control instruction and transmit it to the steering gear. The steering gear steers based on the rudder control instruction to control the rolling aircraft to fly towards the target, thereby completing the present invention.

[0006] Specifically, the purpose of the present invention is to provide a rolling aircraft control method based on a two-loop roll attitude autopilot, in which the roll angle error and roll angle rate are obtained in real time, and the rudder control instruction is further obtained through weighted processing and transmitted to the steering gear. The steering gear steers based on the rudder control instruction to control the rolling aircraft to fly towards the target.

[0007] The rudder deflection control command is obtained by the following formula (1):

[0008]

[0009] Where Δγ represents the roll angle error;

[0010] represents the roll angular rate;

[0011] K A Indicates the roll rudder control command corresponding to the unit roll angle error.

[0012] K g Indicates the roll attitude autopilot damping coefficient.

[0013] Wherein, the roll angle rate Obtained by the following formula (II):

[0014]

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

[0016] k r and T r Each independently expresses the dynamic coefficient;

[0017]

[0018]

[0019] represents the roll damping derivative;

[0020] represents the derivative of the rolling moment produced by the rudder;

[0021] J x Represents the moment of inertia of the aircraft around the longitudinal axis.

[0022] Among them, the k A Obtained by the following formula (III):

[0023]

[0024] Among them, ω CR Indicates the open loop crossing frequency index of the autopilot;

[0025] c δ represents the dynamic coefficient;

[0026]

[0027] Among them, the K g Obtained by the following formula (IV):

[0028]

[0029] Where ζ represents the damping coefficient;

[0030] ω CR Indicates the open loop crossing frequency index of the autopilot;

[0031] c δ and c ω Each independently expresses the dynamic coefficient;

[0032]

[0033]

[0034] The beneficial effects of the present invention include:

[0035] According to the rolling aircraft control method based on the two-loop roll attitude autopilot provided by the present invention, the roll rudder deflection control instruction and the roll attitude autopilot damping coefficient corresponding to the unit roll angle error are constructed based on the autopilot open-loop cross-frequency index and the damping coefficient, so that the roll angle error and the roll angle rate can be more accurately weighted, and more accurate rudder deflection control instructions are provided for subsequent servo steering control, thereby improving the hit accuracy of the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

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

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

[0038] Figure 3 A schematic diagram of the 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 diagram obtained in the comparative example is shown;

[0041] Figure 6 A schematic diagram of the line of sight angle estimation error obtained in the comparative example is shown. DETAILED DESCRIPTION

[0042] The present invention will be further described in detail below through the accompanying drawings and embodiments. Through these descriptions, the characteristics and advantages of the present invention will become more clear and distinct.

[0043] The word "exemplary" is used exclusively herein to mean "serving as an example, embodiment, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Although various aspects of the embodiments are shown in the drawings, the drawings are not necessarily drawn to scale unless otherwise noted.

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

[0045] In a preferred embodiment, the rudder deflection control command is obtained by the following formula (1):

[0046]

[0047] Wherein, Δγ represents the roll angle error; in this application, the roll angle γ of the aircraft is obtained in real time by the inertial navigation unit on the aircraft. c , and then the rolling angle γ c The roll angle error is obtained by subtracting the roll angle γ output by the autopilot.

[0048] represents the roll angular rate;

[0049] K A It represents the forward gain of the roll attitude autopilot, that is, the roll rudder control command corresponding to the unit roll angle error, which is used to control the attitude angle of the aircraft.

[0050] K g It indicates the roll attitude autopilot damping coefficient, which indicates the roll rudder deflection control instruction corresponding to the unit roll angular rate. It is used to improve the aircraft's roll direction damping characteristics and increase the roll damping.

[0051] In this application, by solving the parameter K A and K g , which helps to improve the aircraft's attitude angle and improve the damping characteristics of the aircraft's rolling direction.

[0052] Preferably, the roll angle rate Obtained by the following formula (II):

[0053]

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

[0055] k r and T r Each independently expresses the dynamic coefficient;

[0056]

[0057]

[0058] Represents the roll damping derivative, and its value is determined based on the aircraft's hardware parameters before the aircraft is launched;

[0059] It represents the derivative of the rolling moment generated by the rudder. Its value is determined based on the aircraft's hardware parameters before the aircraft is launched.

[0060] J x Represents the moment of inertia of the aircraft around the longitudinal axis. Its value is determined based on the aircraft's hardware parameters before the aircraft is launched.

[0061] Preferably, the k A Obtained by the following formula (III):

[0062]

[0063] Among them, ω CR It indicates the open-loop crossover frequency index of the autopilot. Its value is determined based on the hardware parameters of the aircraft before the aircraft is launched.

[0064] c δ represents the dynamic coefficient;

[0065]

[0066] Preferably, the Kg is obtained by the following formula (four):

[0067]

[0068] Wherein, ζ represents the damping coefficient, and its value is determined based on the hardware parameters of the aircraft before the aircraft is launched;

[0069] ω CR Indicates the open loop crossing frequency index of the autopilot;

[0070] c δ and c ω Each independently expresses the dynamic coefficient;

[0071]

[0072]

[0073] Example

[0074] For a certain type of rolling aircraft, the values ​​of its power coefficient and related parameters of the aircraft are as follows:

[0075] ω CR =40rad / s

[0076] ζ=0.7

[0077]

[0078]

[0079] J x =0.96kg·m 2

[0080] The aircraft is controlled to intercept a maneuvering target 10 km away. The guidance rate used is the guidance rate provided in the low-cost integrated navigation system and method for high-speed rolling aircraft in Chinese patent 2019103567261. The specific control process is as follows:

[0081] The roll angle error and roll angle rate are obtained in real time, and the rudder control command is further obtained through weighted processing and transmitted to the servo. The servo steers based on the rudder control command to control the rolling aircraft to fly towards the target.

[0082] The rudder deflection control command is obtained by the following formula (1):

[0083]

[0084] Where Δγ represents the roll angle error, which is calculated by c Subtract the roll angle γ output by the autopilot to obtain;

[0085] represents the roll angular rate;

[0086] The roll angle rate Obtained by the following formula (II):

[0087]

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

[0089] k r and T r Each independently expresses the dynamic coefficient;

[0090]

[0091]

[0092] The K A Obtained by the following formula (III):

[0093]

[0094] The K g Obtained by the following formula (IV):

[0095]

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

[0097] Finally, we get K A =0.2199, K g =0.0033.

[0098] The obtained K A and K g Substitute the servo transfer function. In this simulation, the servo transfer function is set to -1. Through simulink simulation, the simulation results are as follows: Figure 1 , Figure 2 and Figure 3 as shown in .

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

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

[0101] The Bode diagram drawn based on the open-loop transfer function is as follows: Figure 3 As shown in , since the servo transfer function is set to -1, it is an ideal state, but in reality the servo is a second-order link. Figure 3 The ω obtained in CR =38.2rad / s, which is consistent with the actual ω CR =40rad / s, and the error is within a reasonable range.

[0102] Comparative Example

[0103] The same aircraft as in Example 1 is used, the same maneuvering target is targeted, and the same guidance rate is used. The control process is as follows:

[0104] The roll angle error and roll angle rate are obtained in real time, and the rudder control command is further obtained through weighted processing and transmitted to the servo. The servo steers based on the rudder control command to control the rolling aircraft to fly towards the target.

[0105] The rudder deflection control command is obtained by the following formula (1):

[0106]

[0107] Where Δγ represents the roll angle error, and the roll angle γ is obtained by c Obtained by subtracting the roll angle γ output by the autopilot;

[0108] represents the roll angular rate;

[0109] The roll angle rate Obtained by the following formula (II):

[0110]

[0111] Where s represents the sign of Laplace transform

[0112] k r and T r Each independently expresses the dynamic coefficient;

[0113]

[0114]

[0115] The K A The value of 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 in;

[0118] pass Figure 4 It can be seen that the curve oscillates severely and gradually diverges, and cannot converge to 1.

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

[0120] The Bode diagram drawn based on the open-loop transfer function is as follows: Figure 6 As shown in CR The corresponding phase margin is less than 0, indicating that the system is unstable.

[0121] It can be seen from the results of the embodiment and the comparative example that the rolling aircraft control method based on the two-loop roll attitude autopilot in the embodiment can ensure that the aircraft flies stably and hits the target more accurately.

[0122] The present invention has been described above in conjunction with preferred embodiments, but these embodiments are only exemplary and serve only as an illustration. On this basis, the present invention may be subjected to a variety of substitutions and improvements, all of which fall within the scope of protection of the present invention.

Claims

1. A rolling aircraft control method based on a two-loop rolling attitude autopilot, characterized in that: In the method, the roll angle error and the roll angle rate are obtained in real time, and the rudder control command is further obtained through weighted processing and transmitted to the servo. The servo steers based on the rudder control command to control the rolling aircraft to fly towards the target.

2. The rolling aircraft control method based on the two-loop rolling attitude autopilot according to claim 1 is characterized in that: The rudder deflection control command is obtained by the following formula (1): Where Δγ represents the roll angle error; represents the roll angular rate; K A Indicates the roll rudder control command corresponding to the unit roll angle error. K g Indicates the roll attitude autopilot damping coefficient.

3. The rolling aircraft control method based on the two-loop rolling attitude autopilot according to claim 2 is characterized in that: The roll angle rate Obtained by the following formula (II): Where s represents the sign of Laplace transform; k r and T r Each independently expresses the dynamic coefficient; represents the roll damping derivative; represents the derivative of the rolling moment produced by the rudder; J x Represents the moment of inertia of the aircraft around the longitudinal axis.

4. The rolling aircraft control method based on the two-loop rolling attitude autopilot according to claim 2 is characterized in that: The K A Obtained by the following formula (III): Among them, ω CR Indicates the open loop crossing frequency index of the autopilot; cδ represents the dynamic coefficient; 5. The rolling aircraft control method based on the two-loop rolling attitude autopilot according to claim 2 is characterized in that: The Kg is obtained by the following formula (four): Where ζ represents the damping coefficient; ω CR Indicates the open loop crossing frequency index of the autopilot; c δ and c ω Each independently expresses the dynamic coefficient;

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