A pseudo angle of attack augmentation and stabilization control method for unmanned aerial vehicles

Abstract

The present invention discloses a method for stabilizing the control of the pseudo angle of attack of an unmanned aerial vehicle, belonging to the technical field of unmanned aerial vehicle flight control. A pitch angle rate signal and a pseudo angle of attack stabilization control strategy are introduced into the elevator control channel to improve the dynamic response characteristics of the unmanned aerial vehicle. The present invention uses the pitch angle rate signal collected by the sensor as feedback, deduces the transfer function from the pitch angle rate to the pseudo angle of attack through relevant theories of flight mechanics, adds pseudo angle of attack feedback to the angular rate stabilization loop, and relaxes the longitudinal static stability through the pseudo angle of attack stabilization; through comparative simulation verification in a six-degree-of-freedom simulation model, the system with the addition of the pseudo angle of attack stabilization loop feedback has better rise time, steady-state error, and overshoot than the system without relaxing the static stability.

Description

Technical Field

[0001] The present invention belongs to the technical field of UAV flight control, and particularly relates to a method for enhancing the stability of the pseudo angle of attack of a UAV. Background Art

[0002] With the increasing diversification and complexity of mission requirements, the requirements for the maneuverability of UAVs are constantly increasing. For example, in some special mission scenarios, UAVs need to quickly change their flight postures and trajectories to adapt to complex environments or complete specific mission objectives. However, the traditional design of UAVs with high static stability restricts the improvement of their maneuverability to a certain extent. For a given UAV flight speed and aerodynamic layout, achieving high maneuverability often requires a large angle of attack. However, in conventional designs, the allowable rudder deflection angle is usually limited, which leads to difficulties in pursuing maneuverability. To solve the above problems, the pseudo angle of attack feedback technology has emerged.

[0003] The pseudo angle of attack feedback technology can achieve the enhancement of the stability of a statically unstable UAV or the reduction of the stability of a statically overly stable UAV and stabilize its control ratio by constructing an angle of attack feedback loop around the UAV body to form an artificial restoring moment. For example, in the design of some missiles, after introducing the pseudo angle of attack feedback, the robustness can be improved while ensuring a high control ratio. When the required angle of attack is large, it can reduce the required steady-state rudder deflection angle and balance the dynamic response requirements to a certain extent.

[0004] In practical applications, it is necessary to more accurately determine the parameters of the pseudo angle of attack feedback loop to better adapt to different flight conditions and mission requirements. At the same time, it is necessary to further study how to better handle problems such as the relationship between rudder deflection angle limitation, flight stability, and maneuverability.

[0005] When the UAV body exhibits low static stability, it is a divergent system, which occurs in the initial power flight stage of some tactical missiles, for example. At this time, without the assistance of an autopilot, it is difficult for the UAV to achieve stable flight. Taking missile control as an example, the transfer function of a rear-controlled missile has right-half plane zeros and belongs to a non-minimum phase system. In the initial stage of control, the acceleration output direction is opposite to the expected direction. It is not until the lift generated by the angle of attack is large enough that the expected acceleration direction will appear, which undoubtedly affects the control accuracy and efficiency. Moreover, traditional autopilots have poor robustness in the face of changes in the body gain. Once the center of gravity and center of pressure positions change, resulting in a large change in the missile body gain, the autopilot design is very likely to fail the robustness assessment.

[0006] In traditional designs, to ensure flight safety and stability, drones are often designed to have high static stability, which makes it difficult for them to quickly change their attitudes and trajectories during flight. For example, in scenarios where agile maneuvers are required to avoid obstacles or quickly track targets, high static stability drones, due to their aerodynamic layouts and control characteristics, cannot quickly respond to commands to adjust their flight states. Existing control methods are difficult to achieve effective relaxed static stability operations on high static stability drones. On the one hand, conventional control strategies lack targeted mechanisms to actively adjust the static stability of drones. During actual flight, when facing complex and changing mission requirements, such as quickly turning in a narrow space or maintaining an accurate flight trajectory under strong airflow interference, existing control methods cannot dynamically reduce the static stability of drones to obtain higher maneuverability; on the other hand, when traditional control methods attempt to adjust static stability, it is easy to cause fluctuations in other flight performances. For example, it may lead to unstable flight attitudes, increased difficulty in controlling flight speeds, etc., thus affecting the overall flight performance and mission execution effectiveness. Therefore, there is an urgent need for a control method that can effectively relax static stability and ensure stable flight performance to meet the requirements of modern drones in diverse mission scenarios. Summary of the Invention

[0007] The present invention provides a method for controlling the augmented stability of the pseudo angle of attack of a drone, introducing a pitch rate signal and a pseudo angle of attack augmented stability control strategy in the elevator control channel to improve the dynamic response characteristics of the drone.

[0008] To achieve the above objectives, the present invention adopts the following technical solutions:

[0009] A method for controlling the augmented stability of the pseudo angle of attack of a drone mainly introduces a pitch rate signal in the elevator control channel to improve the dynamic response characteristics of the drone, specifically including the following steps:

[0010] Taking the pitch rate signal collected by the sensor as feedback, deriving the transfer function from the pitch rate to the pseudo angle of attack through relevant flight mechanics theories, adding a pseudo angle of attack feedback in the angular rate augmented stability loop, and relaxing the longitudinal static stability through the pseudo angle of attack augmented stability.

[0011] The expression of the transfer function from the pitch rate to the pseudo angle of attack is: ,

[0012] Where is the pitch rate signal measured by the sensor, is the complex variable in the Laplace transform, is the pseudo angle of attack, is the dynamic derivative: the longitudinal force caused by the change in the angle of attack.

[0013] The augmented stability control law for the pitch channel based on classical feedback control is:

[0014] ,

[0015] Among them, is the elevator command of the stability augmentation loop, is the elevator command of the pseudo angle of attack stability augmentation loop, is the elevator command of the original angular rate stability augmentation loop, is the pseudo angle of attack stability augmentation control gain, is the pitch angular rate stability augmentation control gain.

[0016] Beneficial effects: The present invention provides a method for pseudo angle of attack stability augmentation control of an unmanned aerial vehicle. By introducing the pitch angular rate signal into the elevator control channel, the dynamic response characteristics of the unmanned aerial vehicle are improved; the pitch angular rate signal collected by the sensor is used as feedback, and the transfer function from the pitch angular rate to the pseudo angle of attack is derived through relevant flight mechanics theories. A pseudo angle of attack feedback is added to the angular rate stability augmentation loop, and the longitudinal static stability is relaxed through pseudo angle of attack stability augmentation. At an altitude of 4 km, an indicated airspeed of 130 m / s, and a flight simulation is carried out under the condition of consuming half of the fuel. A six-degree-of-freedom flight simulation is performed, and a pitch angle step response is given. It can be seen from the time-domain simulation that for the system with the pseudo angle of attack stability augmentation loop feedback added in the present invention, the rise time, steady-state error, and overshoot are all better than those of the system without relaxed static stability. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 is the pseudo angle of attack calculation block diagram in the embodiment of the present invention;

[0018] Figure 2 is the pseudo angle of attack stability augmentation loop block diagram in the embodiment of the present invention;

[0019] Figure 3 is the angle of attack - elevator root locus diagram in the embodiment of the present invention;

[0020] Figure 4 is the attitude response curve diagram in the embodiment of the present invention. DETAILED DESCRIPTION OF THE INVENTION

[0021] The present invention will be described in detail below with reference to the drawings and specific embodiments:

[0022] In the initial stage of the longitudinal movement of the unmanned aerial vehicle, the short period dominates. In order to improve the damping characteristics and stability of the aircraft, the pitch channel stability augmentation system generally uses pitch angular rate signal feedback. However, this loop will reduce the static maneuverability of the aircraft. The state space equation of the short period mode is obtained by linearizing the non-linear mathematical model:

[0023] ,

[0024] ,

[0025] ,

[0026] Select a trim point as: , , , , , where , where represents the flight speed, represents the flight altitude, represents the angle of attack, represents the pitch rate, represents the elevator command, represents the throttle command.

[0027] Short-period natural frequency , , the natural frequency is relatively high. The root locus from the angle of attack to the elevator is as shown in Figure 3 . As the open-loop gain increases slightly, the closed-loop poles fall on the real axis, becoming an overdamped system, and the frequency is too large, reflecting that the system has a tendency to return to stability too quickly, and the system stability is too large. Therefore, the static stability needs to be relaxed.

[0028] The following provides a pseudo angle-of-attack augmentation control method for an unmanned aerial vehicle with relaxed static stability, which specifically includes the following steps:

[0029] Give the natural form of the longitudinal small perturbation equation:

[0030] ,

[0031] where is the mass of the aircraft, is the engine installation angle, is the angle of attack under the reference motion, is the engine thrust under the reference motion, is the flight path angle under the reference motion, is the speed under the reference motion, is the change in the angle of attack, is the change in the pitch angle, is the change in the flight path angle under the reference motion, is the change in altitude, is the change in speed, is the change in the elevator, is the change in the throttle lever, is the drag caused by the change in the angle of attack, is the drag caused by the change in speed, is the drag caused by the change in altitude, is the drag caused by elevator deflection, is the lift caused by the change in angle of attack, is the lift caused by the rate of change of angle of attack, is the lift caused by the change in speed, is the lift caused by the change in pitch rate, is the lift caused by the change in altitude, is the lift caused by elevator deflection, is the pitching moment caused by the change in angle of attack, is the pitching moment caused by the rate of change of angle of attack, is the pitching moment caused by the change in speed, is the pitching moment caused by the change in pitch rate, is the pitching moment caused by the change in altitude, is the pitching moment caused by elevator deflection, is the engine thrust caused by the change in speed, is the engine thrust caused by the change in altitude, is the engine thrust caused by the change in throttle lever.

[0032] For the analysis of aircraft stability and controllability, generally, the influence of altitude change on external forces and external moments can be ignored, and it is assumed that the reference motion is horizontal steady straight flight. After simplification, we can get:

[0033] ,

[0034] where the state vector , the control vector ;

[0035] Omitting the tangential force equation, we get an approximate equation set that mainly reflects the initial stage of disturbance recovery:

[0036] .

[0037] Taking the pitch rate signal collected by the sensor as feedback, the transfer function from pitch rate to pseudo angle of attack is deduced through the above flight mechanics related theories. Adding pseudo angle of attack feedback to the angular rate stability augmentation loop relaxes the longitudinal static stability.

[0038] The expression of the transfer function from pitch rate to pseudo angle of attack is: ,

[0039] where is the pitch rate signal measured by the sensor, is the complex variable in the Laplace transform, is the pseudo angle of attack, is the dynamic derivative: the longitudinal force caused by the change in angle of attack.

[0040] The transfer function from pitch rate to pseudo angle of attack, after discretization in the Simulink module, has the following formula:

[0041] ,

[0042] The calculation block for the pseudo angle of attack can be obtained as Figure 1 shown.

[0043] As Figure 2 shown, taking the pitch rate acquired by the pitch sensor as feedback, the pseudo angle of attack is obtained through the transfer function from pitch rate to pseudo angle of attack. Then, through the dynamic equation and the proportional term, the transfer function from the pseudo angle of attack to the elevator can be obtained: , as the compensation part for pseudo angle of attack stability augmentation / relaxed static stability, plus the original pitch rate stability augmentation feedback, the pitch channel stability augmentation control law based on classical feedback control can be obtained: ,

[0044] Among them, is the elevator command for the pseudo angle of attack stability augmentation loop, is the control gain for pseudo angle of attack stability augmentation, is the dynamic derivative: the pitching moment caused by the change in angle of attack, is the dynamic pressure acquired by the air data computer, is the wing area, is the mean geometric chord length, is the moment of inertia about the body y-axis; is the elevator command for the stability augmentation loop, is the elevator command for the original angle rate stability augmentation loop, is the control gain for pitch rate stability augmentation.

[0045] In the six-degree-of-freedom simulation model, a comparative simulation verification is carried out. In the flight condition of an altitude of 4 km, an indicated airspeed of 130 m / s, and half of the fuel consumed, a six-degree-of-freedom flight simulation is performed. Given a pitch angle step response, the time-domain simulation results are as Figure 4 shown. It can be seen that the rise time, steady-state error, and overshoot of the system with the feedback of the pseudo angle of attack stability augmentation loop are all better than those of the system without relaxed static stability, indicating that the above method effectively improves the dynamic response characteristics of the UAV.

[0046] The above are only the preferred embodiments of the present invention. It should be noted that for those of ordinary skill in the art, without departing from the principle of the present invention, several improvements and refinements can be made, and these improvements and refinements should also be regarded as the protection scope of the present invention.

Claims

1. A pseudo angle of attack stabilization control method for an unmanned aerial vehicle, characterized in that: The following steps are involved: This gives the natural form of the longitudinal small perturbation equation: Where m is the mass of the aircraft, is the engine installation angle, α * is the angle of attack under reference motion, T * is the engine thrust under reference motion, γ * is the track inclination angle under reference motion, V * is the velocity under the reference motion, Δα is the change in angle of attack, Δθ is the change in pitch angle, Δγ * is the change in track inclination under the reference motion, ΔH is the change in height, ΔV is the change in speed, and Δδ e is the elevator change, Δδ T is the change of throttle lever, D α is the drag caused by the change in angle of attack, D V is the resistance caused by speed change, D H is the resistance caused by the height change, is the drag caused by elevator deflection, L α is the lift caused by the change in angle of attack, is the lift caused by the rate of change of angle of attack, L V is the lift caused by the speed change, L q is the lift caused by the change of pitch rate, L H is the lift caused by the altitude change, is the lift caused by elevator deflection, M α is the pitching moment caused by the change in angle of attack, is the pitch moment caused by the rate of change of angle of attack, M V is the pitch moment caused by speed change, M q is the pitch moment caused by the change of pitch angle rate, M H is the pitching moment caused by the altitude change, is the pitching moment caused by the elevator deflection, T V is the engine thrust caused by speed change, T H is the engine thrust caused by the altitude change, The engine thrust caused by the change of throttle lever; Ignoring the influence of altitude change on external forces and torques, and assuming the reference motion to be horizontal steady straight flight, the approximate equation for the initial stage of main reaction disturbance recovery is obtained: Where Z α is the longitudinal force caused by the change in angle of attack; The pitch rate signal collected by the sensor is used as feedback, and pseudo angle of attack feedback is added to the angle rate stabilization loop. The longitudinal static stability is relaxed by pseudo angle of attack stabilization. The transfer function from pitch rate to pseudo angle of attack is derived through the above flight mechanics theory: Where q is the pitch rate signal measured by the sensor, s is the complex variable in Laplace transform, α is the pseudo angle of attack, and the pseudo angle of attack is obtained through the transfer function from pitch rate to pseudo angle of attack; The transfer function from the pseudo angle of attack to the elevator is used as the compensation part of the pseudo angle of attack stabilization / relaxed stabilization, and combined with the original pitch angle rate stabilization feedback, the pitch channel stabilization control law based on classical feedback control is obtained: in, is the elevator command for the stability augmentation loop, is the elevator command of the pseudo angle of attack stabilization loop, is the elevator command of the original angular rate stabilization loop, K α is the pseudo angle of attack stabilization control gain, K q Pitch rate stabilization control gain.

2. The pseudo angle of attack stabilization control method for an unmanned aerial vehicle according to claim 1, characterized in that: For the analysis of aircraft stability and maneuverability, the influence of altitude change on external forces and moments is ignored, and the reference motion is assumed to be horizontal steady straight flight. The perturbation equation is simplified to: Where state vector x=[ΔV Δa Δq Δθ] T , manipulate vector u=[Ds e Dd T ] T 。 3. The pseudo angle of attack stabilization control method for an unmanned aerial vehicle according to claim 1, characterized in that: The transfer function from pitch rate to pseudo angle of attack is discretized in the simulink module as follows: [-α(k)·Z α +q(k)]·ΔT+α(k-1)=α(k) Get the calculation frame of the pseudo angle of attack.

4. The pseudo angle of attack stabilization control method for an unmanned aerial vehicle according to claim 1 or 3, characterized in that: The transfer function of the pseudo angle of attack to the elevator is: in, is the elevator command of the pseudo angle of attack stabilization loop, K α is the pseudo angle of attack stabilization control gain, M α The pitching moment caused by the change in angle of attack, is the dynamic pressure of the atmospheric machine, S is the wing area, c is the average geometric chord length, I y is the moment of inertia about the y-axis of the machine body.

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