Direct lift active disturbance rejection control method suitable for flying wing layout

Abstract

The invention belongs to the technical field of glide control, and relates to a direct lift active disturbance rejection control method suitable for flying wing layout. According to the method, a trailing edge flap is introduced to serve as a main lift control surface, meanwhile, a constant attack angle is kept through a lifting aileron so as to balance part of pitching moment generated by the trailing edge flap, and therefore the attitude of the aircraft is stabilized, and the constant speed is maintained. In the actual flight process, external disturbance is inevitable, so that an original glide control strategy is improved and optimized in combination with an active disturbance rejection control method. Compared with a traditional glide control method, the flying wing layout direct lift active-disturbance-rejection control method provided by the invention has remarkable anti-interference capability and excellent glide control performance, and shows a wide application prospect.

Description

Technical Field

[0001] The invention belongs to the technical field of glide control and relates to a direct lift self-disturbance rejection control method suitable for a flying wing layout. Background Art

[0002] The tailless flying wing layout eliminates the vertical tail and horizontal tail, and integrates the fuselage and wings, greatly simplifying the fuselage structure, with outstanding advantages such as light structure weight, low flight resistance, and good stealth performance. For the conventional layout, the rudder on the vertical tail controls the flight direction of the aircraft, the elevator on the horizontal tail controls the pitch attitude of the aircraft, and the aileron on the wing controls the roll of the aircraft. For the tailless flying wing layout, the functions of the rudder, elevator and aileron must be integrated to achieve the three-in-one control of direction, pitch and roll. Therefore, each rudder action may have the control effect of changing the direction, pitch and roll at the same time. There is a coupling problem of the control rudder, which requires decoupling and high-precision control of the rudder according to the control intention. Therefore, high-precision control of flying wing layout aircraft is a research direction with great practical significance.

[0003] In the traditional control mode, when a glide path deviation occurs, the aircraft needs to generate a control torque through the control surface control command to achieve a change in attitude. The change in attitude will cause a change in aerodynamic force, thereby correcting the glide path deviation. This conventional torque control strategy can only indirectly adjust the trajectory by controlling the change in the control surface, torque, attitude, and force. There is a certain time lag, which is not conducive to the rapid response of control. In order to improve the quality of glide control, direct lift began to be applied in glide control. By introducing direct lift control, the relevant parameters can be directly adjusted to adjust the flight trajectory. Compared with the traditional control law, the control law introducing direct lift realizes the decoupling of trajectory and attitude control, but this solution is mainly implemented in fixed-wing aircraft. For tailless flying wing layout aircraft, the introduction of direct lift technology is a key technology.

[0004] During the flight of a flying wing aircraft, it will be disturbed by many air currents. How to dynamically and effectively control the large-scale uncertainties with strong nonlinearity, strong coupling, fast time-varying characteristics in such a complex flight environment is a serious problem faced in flight control. The direct lift anti-disturbance control method for flying wing configuration is an effective control method. It can observe and estimate in real time through the state expansion observer and eliminate the "total disturbance" through compensation, so that the aircraft control system full of disturbances, uncertainties and nonlinearities is transformed into an ideal standard integral series form, thereby achieving better control effect.

[0005] In published research, the direct lift glide control method has been used in many occasions, but it has not been applied to flying wing tailless aircraft. However, its feasibility has been verified. Wang Yanxiong (Research on Take-off and Landing Control Technology of Flying Wing UAV [D]. Northwestern Polytechnical University, 2017.) proposed a composite control scheme that combines the control of the additional lift of the rudder surface with the aerodynamic control of the fuselage. Although the trajectory control is not completely handed over to the additional lift, it can reduce the height loss of the flying wing aircraft when encountering a micro-downdraft compared to the traditional control method. Wu Wenhai (Analysis of MAGIC CARPET Landing Technology [J]. Systems Engineering and Electronic Technology, 2018, 40(09): 2079-2091.) and other scholars analyzed the key technologies of direct lift and conceived a comprehensive direct side force technology. Scholars such as Zhang Zhibing (A Review of Automatic Landing Guidance and Control of Carrier-based Aircraft [J]. Journal of Nanjing University of Aeronautics and Astronautics, 2018, 50(06): 734-744. DOI: 10.16356 / j.1005-2615.2018.06.002.) and Zhu Yulian (Research on "Magic Carpet" Landing Technology of Carrier-based Aircraft [D]. Nanjing University of Aeronautics and Astronautics, 2020. DOI: 10.27239 / d.cnki.gnhhu.2020.000195.) clarified the configuration of the direct lift controller, adopted a composite control configuration with flaps as the main and elevators as the auxiliary, and achieved trajectory control through direct lift flap control surfaces. Deng Jinlai (DOI:10.19555 / j.cnki.1673-4599.2020.02.002.) and other scholars analyzed the direct force control mode, preliminarily conceived and designed aircraft PID control laws under three direct force modes, and conducted accuracy comparison tests. In the field of control algorithms, Luo Fei (DOI: 10.13645 / j.cnki.fd20201113.010.), Song Liting (Cascaded preset performance dynamic inversion decoupling direct lift landing control [J]. Journal of Harbin Institute of Technology, 2023, 55(12): 42-53.) and other scholars used dynamic inversion, improved dynamic inversion, incremental dynamic inversion and other algorithms to optimize the traditional control law, and Wu Qilong (Longitudinal carrier-based aircraft direct lift fully automatic landing control based on linear anti-disturbance control [J]. Journal of Intelligent Systems, 2024, 19(01): 142-152.) and others used anti-disturbance control algorithms to improve the traditional control law. In addition, Liu Rendi and Jiang Ju (Direct lift control technology for carrier-based aircraft landing based on reinforcement learning [J / OL]. Journal of Beijing University of Aeronautics and Astronautics, 1-17 [2025-03-04]. https: / / doi.org / 10.13700 / j.bh.1001-5965.2023.0403.) also adopted intelligent optimization methods such as neural networks, but there are few studies that apply the anti-interference control method based on direct lift to flying wing layout aircraft. Summary of the invention

[0006] In order to solve the coupling problem of attitude and trajectory in traditional glide control, the present invention adopts a glide control method based on direct lift mode. The method firstly introduces the trailing edge flap as the direct lift control surface, and the elevon keeps the angle of attack constant and balances part of the pitch moment generated by the trailing edge flap, so as to stabilize the attitude of the aircraft, and adopts throttle control to maintain the stability of the flight speed. On this basis, in view of the uncertainty in the glide process of the flying wing layout aircraft, the self-anti-disturbance control method is introduced into the angle of attack holding loop and the power compensation loop for optimization design, which improves the anti-disturbance capability and control quality. The simulation shows that the method has better robustness and control accuracy in the presence of disturbance.

[0007] The technical solution of the present invention:

[0008] A direct lift anti-disturbance control method suitable for flying wing layout, including flying wing layout aircraft dynamics modeling, direct lift-based glide control design and anti-disturbance control method design. The details are as follows:

[0009] (1) Dynamics modeling of flying wing aircraft

[0010] During the aircraft's descent, the dynamic model of the aircraft's basic longitudinal motion is:

[0011]

[0012] Where: V is speed, T is thrust, α is angle of attack, D is drag, m is mass, g is acceleration due to gravity, γ is track angle, L is lift, I y is the moment of inertia, q is the pitch angular velocity, M is the pitch moment, x g is the forward displacement distance, H is the flight altitude, and θ is the pitch angle.

[0013]

[0014] Where: D is the drag, L is the lift, M is the pitching moment, C D is the drag coefficient, C L is the lift coefficient, C m is the pitching moment coefficient, ρ is the atmospheric density, S is the wing area, and c is the mean aerodynamic chord length.

[0015]

[0016] Where: PF is the trailing edge flap, δ ELE For the elevons, is the zero deflection resistance derivative, is the drag derivative of the angle of attack, is the elevon drag derivative, is the drag derivative of the trailing edge flap, is the zero deflection lift derivative, is the lift derivative of the angle of attack, is the lift derivative of the elevon, is the lift derivative of the trailing edge flap, C m0 is the zero-bias pitch moment derivative, C mα is the pitching moment derivative of the angle of attack, is the trailing edge flap pitch moment derivative, is the elevon pitching moment derivative, and H is the height.

[0017] (2) Descent control design based on direct lift

[0018] Overall idea: The direct lift control mode includes two modes: flight trajectory rate mode and flight trajectory increment mode. The flight trajectory rate mode controls the descent trajectory by adjusting the trajectory angle rate, thereby quickly and accurately adjusting the glide path and improving the dynamic performance of the glide path control. The flight trajectory increment mode automatically maintains the ideal glide angle by adjusting the increment of the trajectory angle, thereby changing the altitude to ensure the steady-state accuracy of the glide path control. However, both schemes can achieve stable glide control of the aircraft.

[0019] The present invention studies the flight trajectory rate mode and divides the design into an altitude loop, a direct lift control loop, an attitude loop based on a constant angle of attack, and a power compensation loop based on a constant speed. The classical PID control law is used in the altitude loop and the direct lift control loop, and the anti-disturbance control law is used for the attitude loop based on a constant angle of attack and the power compensation loop based on a constant speed. The control laws of the altitude loop and the direct lift control loop are as follows:

[0020] The trailing edge flaps are selected as the control surfaces for direct lift to control the track angle. In the process of solving the control command, the track angle rate command can be calculated according to the deviation of the gliding process. By controlling the change of the trailing edge flaps, the track angle rate can be directly controlled to eliminate the deviation of the track angle.

[0021]

[0022] Among them, δ PF0 is the trailing edge flap trim value, is the actual value of the track angular velocity, is the reference track angular rate, K p,H is the proportional control parameter of the trajectory loop controller, K i,H is the integral control parameter of the trajectory loop controller, K d,H is the differential control parameter of the trajectory loop controller, is the proportional control parameter of the direct lift loop controller, is the integral control parameter of the direct lift loop controller, H is the actual height, H c is the height reference instruction, and s is the complex variable of Laplace transform.

[0023] (3) Design of Active Disturbance Rejection Control Method

[0024] The anti-disturbance control method is adopted for the attitude loop based on constant angle of attack and the power compensation loop based on constant speed, as follows:

[0025] (3.1) Attitude loop based on constant angle of attack

[0026] In the direct lift control mode, the angle of attack needs to be kept constant, and the angle of attack holding circuit controls the angle of attack by introducing the elevon auxiliary control surface.

[0027] According to the aircraft dynamics formula, the following relationship is obtained:

[0028]

[0029] The derivative relationship of the angle of attack is as follows:

[0030]

[0031] in, is the first-order derivative of the pitch angle, is the first derivative of the angle of attack.

[0032] Linearizing the angle of attack at the reference point into the following equation:

[0033]

[0034] Substitute it into the dynamic model and we get:

[0035]

[0036] After sorting, we can get the following relationship:

[0037]

[0038] Among them, in order to comprehensively consider the first-order inertial link transfer function of the elevator aileron actuator, it is expanded and rewritten as:

[0039]

[0040] in, is the second-order derivative of the angle of attack, f 1 and b 1 Represents a relationship for the middle term.

[0041] Expanding the total disturbance in the angle of attack holding loop into a new state variable, the system is further expressed as:

[0042]

[0043] Among them, α 1 , α 2 , α 3 To expand variables, is the first-order derivative of the corresponding expanded variable, f 1 The first derivative of .

[0044] The extended state observer designed for the above system is as follows:

[0045]

[0046] Where: It is for α 1 An accurate estimate of Yes An accurate estimate of is an estimate of the total disturbance; β 11 , β 12 and β 13 is the extended state observer gain of the angle of attack holding loop, β 11 =3ω 1 , ω 1 For bandwidth.

[0047] The following control law is obtained by introducing the tracking differentiator and state error feedback:

[0048]

[0049] Where: e α1 is the angle of attack error signal, e α2 is the angle of attack differential error signal, u 1 is the error feedback control output, α c is the desired tracking signal at angle of attack, is the desired differential signal of the angle of attack; is the proportional control parameter of the angle of attack holding loop, is the differential control parameter of the angle of attack holding loop.

[0050] Since the deflection of the trailing edge flap will produce some pitching moment, it is necessary to introduce the deflection of the elevon for compensation and decoupling, as shown in the following formula:

[0051]

[0052] Where: K is the torque decoupling coefficient.

[0053] (3.2) Power compensation circuit based on constant speed

[0054] The velocity equation is linearized as follows:

[0055]

[0056] in, is the first-order derivative of velocity.

[0057] The aircraft's thrust engine model is related to the throttle:

[0058]

[0059] Substitute it into the dynamic model and we get:

[0060]

[0061] After rearranging, we can get the following formula:

[0062]

[0063] Since the transfer function of the throttle lever actuator is also a first-order inertia link, the throttle lever transfer function is rewritten by comprehensively considering it:

[0064]

[0065] in, is the second-order derivative of velocity, f 2 and b 2 Represents a relationship for the middle term.

[0066] Consider the total disturbance in the dynamic compensation loop as an expanded state variable:

[0067]

[0068] Among them, V 1 、V 2 、V 3 To expand variables, is the first-order derivative of the corresponding expanded variable, f 2 The first derivative of .

[0069] The third-order active disturbance rejection control law is established as follows:

[0070]

[0071] Where: is the speed error signal, is the speed differential error signal, is the speed estimation error signal, V c and is the desired tracking and differentiation signal of the velocity; Yes V 1 estimates; yes estimates; is an estimate of the total disturbance; β 21 , β 22 and β 23 is the extended state observer gain of the dynamic compensation loop; β 21 =3ω 2 , ω 2 is the bandwidth; δ T is the throttle command; is the proportional control parameter of the power compensation loop, It is the differential control parameter of the power compensation loop.

[0072] In summary, the present invention adopts an anti-disturbance control method for the angle of attack holding link and the power compensation link, and adopts a traditional control method for the altitude loop and the direct lift control loop.

[0073] Beneficial effects of the present invention:

[0074] Aiming at the attitude trajectory coupling problem during the descent of a flying wing aircraft, the present invention designs a direct lift control loop, an angle of attack holding control loop and a power compensation control loop respectively, and verifies the control system in a disturbance environment. On this basis, aiming at the interference uncertainty problem existing in the descent process, the direct lift control system is designed using the anti-disturbance control technology. This method can automatically observe and compensate for external interference, which is beneficial to the rapid and accurate correction of the trajectory during the aircraft descent mission. The simulation results show that the proposed direct lift anti-disturbance control method for flying wing layout exhibits good anti-interference ability and strong trajectory tracking ability when facing interference, providing an effective solution for improving the safety and control accuracy of the descent process of flying wing aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0075] Figure 1 It is a block diagram of the direct lift anti-disturbance control method for flying wing configuration;

[0076] Figure 2 This is a diagram of the control surface configuration of a flying wing aircraft;

[0077] Figure 3 It is the block diagram of the flight trajectory rate modal control structure;

[0078] Figure 4 It is the structure diagram of the second-order active disturbance rejection controller;

[0079] Figure 5 This is a block diagram of the active disturbance rejection control method based on direct lift;

[0080] Figure 6a It is the simulation result of height comparison of the self-disturbance rejection glide control based on direct lift;

[0081] Figure 6b It is the simulation result of speed comparison of the self-disturbance rejection glide control based on direct lift;

[0082] Figure 6c This is the comparison simulation result of the angle of attack of the self-disturbance rejection glide control based on direct lift;

[0083] Figure 6d It is the simulation result of pitch angle comparison based on direct lift self-disturbance anti-gliding control;

[0084] Figure 6e It is the simulation result of the track angle rate comparison based on the direct lift self-disturbance anti-gliding control;

[0085] Figure 6f It is the simulation result of the track angle comparison based on the direct lift self-disturbance anti-gliding control;

[0086] Figure 6g It is the comparison simulation result of the trailing edge flap of the direct lift self-disturbance anti-gliding control;

[0087] Figure 6h It is the simulation result of the comparison of elevons based on the direct lift self-disturbance anti-gliding control;

[0088] Figure 7a It is the simulation result of the speed comparison of the self-disturbance rejection glide control under the angle of attack disturbance;

[0089] Figure 7b It is the simulation result of the angle of attack comparison of the automatic anti-disturbance glide control under the angle of attack disturbance;

[0090] Figure 7c It is the simulation result of the track angle rate comparison of the anti-disturbance glide path control under the angle of attack disturbance;

[0091] Figure 7d It is the simulation result of the track angle comparison of the anti-disturbance glide path control under angle of attack interference. DETAILED DESCRIPTION

[0092] The specific implementation of the present invention is further described below in conjunction with the accompanying drawings and technical solutions.

[0093] The overall flow chart of the direct lift anti-disturbance control method for flying wing configuration is as follows: Figure 1 As shown, first according to Figure 2 The dynamic modeling of the flying wing layout aircraft control surface configuration is carried out, and then the glide control structure based on direct lift is analyzed and designed. Finally, the anti-disturbance control method is introduced and designed for the angle of attack holding loop and the power compensation loop to improve the control quality. Figure 3 The control algorithm process is used to design direct lift and glide control, and then combined with Figure 4 The second-order auto-disturbance rejection principle block diagram is shown in the figure. Figure 5The direct lift anti-disturbance control method shown in the figure designs the anti-disturbance control method for the power compensation loop and the angle of attack holding loop, and finally compares it with the conventional method through simulation analysis.

[0094] (1) Input the initial state and give the target state

[0095] This embodiment takes a flying wing tailless aircraft as the research object, takes the glide path control as the working condition for control design, takes the flight trajectory rate mode based on the direct lift mode as the basic controller, and performs an anti-disturbance control comparative design on it. The initial conditions of the simulation are shown in Table 1, and the simulation time is 89s:

[0096] Table 1 Simulation parameters

[0097]

[0098] (2) Establishment of dynamic model

[0099] The forces acting on an aircraft in flight include its own gravity mg, engine thrust T and aerodynamic force.

[0100] The engine thrust is related to the throttle opening, flight Mach number and flight altitude. When it is 0, the engine thrust is:

[0101] T=f(δ T ,V,H)

[0102] Where m is the weight of the aircraft, g is the acceleration due to gravity, and δ T is the throttle opening, V is the flight speed, H is the flight altitude, and T is the engine thrust.

[0103] Convert each aerodynamic coefficient to the body axis to obtain the aerodynamic force expressed by the aircraft on the body axis:

[0104]

[0105] Where L is lift, D is drag, Y is side force, ρ is air density, S is wing reference area, C L is the lift coefficient, C D is the drag coefficient, C Y is the side force coefficient.

[0106] The aerodynamic coefficient is expressed as:

[0107]

[0108] in, is the zero deflection lift derivative, is the lift derivative of the angle of attack, is the lift derivative of the elevon, is the lift derivative of the trailing edge flap, is the lift derivative of the drag rudder, is the zero deflection resistance derivative, is the drag derivative of the angle of attack, is the elevon drag derivative, is the drag derivative of the trailing edge flap, is the rudder drag derivative, is the zero lateral force derivative, is the lateral force derivative of the drag rudder, is the lateral force derivative of the rolling angular velocity, is the lateral force derivative of the yaw angular velocity, is the elevon side force derivative, is the side force derivative of the trailing edge flap, α is the angle of attack, β is the sideslip angle, and are the roll angular velocity and yaw angular velocity of dimension 1, p is the roll angular velocity, r is the yaw angular velocity, b is the wing span, and the aircraft's trailing edge flaps and elevons are used as direct lift mechanisms. PF is the trailing edge flap input, δ ELE is the elevon input, δ SSD Input for the drag rudder.

[0109] By converting each aerodynamic moment coefficient to the body axis, we can obtain the rolling moment, pitching moment and yaw moment of the aircraft around the body axis in the body axis system:

[0110]

[0111] in, is the rolling moment, M is the pitching moment, and N is the yaw moment. c is the average aerodynamic chord length, C l is the rolling moment coefficient, C m is the pitching moment coefficient, C n is the yaw moment coefficient.

[0112] The aerodynamic moment coefficient is expressed as:

[0113]

[0114] In the formula, C lβ is the rolling moment derivative of the sideslip angle, is the rolling moment derivative of the drag rudder, is the rolling moment derivative of the trailing edge flap, is the elevon rolling moment derivative, is the rolling moment derivative of the rolling angular velocity, is the yaw angular velocity roll moment derivative, C m0 is the zero-bias pitch moment derivative, C mαis the pitching moment derivative of the angle of attack, is the drag rudder pitching moment derivative, is the trailing edge flap pitching moment derivative, is the elevon pitching moment derivative, C nβ is the yaw moment derivative of the sideslip angle, is the yaw moment derivative of the drag rudder, is the yaw moment derivative of the trailing edge flap, is the elevon yaw moment derivative, is the yaw moment derivative of the rolling angular velocity, is the yaw moment derivative of the yaw angular velocity.

[0115] The dynamics formula for aircraft dynamics modeling is as follows:

[0116]

[0117] Among them, F x is the component of the thrust and aerodynamic force on the x-axis of the fuselage, F y is the component of the thrust and aerodynamic force on the y-axis of the aircraft, F z It is the component of the thrust and aerodynamic force on the z-axis of the fuselage. u is the forward velocity component in the fuselage coordinate system, w is the vertical velocity component in the fuselage coordinate system, and v is the lateral velocity component in the fuselage coordinate system. is the forward acceleration component in the body coordinate system, is the vertical acceleration component in the body coordinate system, is the lateral acceleration component in the body coordinate system. p is the roll angular velocity, q is the pitch angular velocity, and r is the yaw angular velocity. is the roll angular acceleration, is the pitch angular acceleration, is the yaw acceleration. φ is the roll angle, θ is the pitch angle, and ψ is the yaw angle. is the first derivative of the roll angle, is the first derivative of the pitch angle, is the first derivative of the yaw angle. is the first derivative of the forward distance, is the first derivative of the lateral distance, is the first derivative of the height. (I x ,I y ,I z ) is the moment of inertia of the three axes, (I xy ,I zy ,I xz ) is the product of inertia,

[0118] During the descent of an aircraft, longitudinal motion is a key factor affecting the control accuracy and safety of descent. Therefore, the present invention mainly focuses on the modeling and control of longitudinal motion. The American coordinate system is selected to establish a longitudinal dynamic model. The dynamic model describing the basic longitudinal motion of the aircraft is:

[0119]

[0120] Where: V is speed, T is thrust, α is angle of attack, D is drag, m is mass, g is acceleration due to gravity, γ is track angle, L is lift, I y is the moment of inertia, q is the pitch angular velocity, M is the pitch moment, x g is the forward displacement distance, H is the flight altitude, and θ is the pitch angle.

[0121]

[0122] Where: D is the drag, L is the lift, M is the pitching moment, C D is the drag coefficient, C L is the lift coefficient, C m is the pitching moment coefficient, ρ is the atmospheric density, S is the wing area, and c is the mean aerodynamic chord length.

[0123]

[0124] Where: PF is the trailing edge flap, δ ELE For the elevons, is the zero deflection resistance derivative, is the drag derivative of the angle of attack, is the elevon drag derivative, is the drag derivative of the trailing edge flap, is the zero deflection lift derivative, is the lift derivative of the angle of attack, is the lift derivative of the elevon, is the lift derivative of the trailing edge flap, C m0 is the zero-bias pitch moment derivative, C mα is the pitching moment derivative of the angle of attack, is the trailing edge flap pitch moment derivative, is the pitch moment derivative of the elevator, H is the height (the drag rudder of the simple longitudinal glide process δ SSD No deflection, simplified aerodynamic model).

[0125] (3) Simulation of glide control based on direct lift

[0126] In order to compare the active disturbance rejection controller of the present invention, PID control laws are first designed for the altitude loop, direct lift loop, attitude loop based on constant angle of attack, and power compensation loop based on constant speed, and baseline simulation is performed. The specific design process is shown below.

[0127] (3.1) Overall design concept of descent control based on direct lift

[0128] The study is based on the flight trajectory rate mode in the direct lift control mode. The track angle is precisely controlled by adjusting the track angle rate through deflecting the trailing edge flaps, thereby changing the altitude. The elevon deflection is used to keep the angle of attack constant and the pitch moment balanced. In addition, a power compensator based on constant speed is used to keep the speed constant through the throttle.

[0129] (3.2) Direct lift control structure based on flight trajectory rate mode

[0130] The trailing edge flaps are selected as the control surfaces for direct lift to control the track angle. In the process of solving the control command, the track angle rate command can be calculated according to the deviation of the gliding process. By controlling the change of the trailing edge flaps, the track angle rate can be directly controlled to eliminate the deviation of the track angle.

[0131]

[0132] Among them, δ PF0 is the trailing edge flap trim value, is the actual value of the track angular velocity, is the reference track angular rate, K p,H is the proportional control parameter of the trajectory loop controller, K i,H is the integral control parameter of the trajectory loop controller, K d,H is the differential control parameter of the trajectory loop controller, is the proportional control parameter of the direct lift loop controller, is the integral control parameter of the direct lift loop controller, H is the actual height, H c is the height reference instruction, and s is the complex variable of Laplace transform.

[0133] (2.2) Attitude loop control structure based on constant angle of attack

[0134] Since the direct lift control structure requires the aircraft's angle of attack to be constant, under this condition, the change in the aircraft's pitch angle is the same as the change in the track angle. However, when correcting the aircraft's trajectory, it is not necessary to rely on changing the pitch angle. Therefore, the elevons are used to control the angle of attack to keep the angle of attack constant.

[0135]

[0136] Among them, δELE0 is the trim value of the elevon, α is the actual value of the angle of attack, α c is the angle of attack reference command, K is the torque decoupling coefficient, K q is the damping coefficient, K p,α is the proportional control parameter of the angle of attack holding loop, K i,α It is the integral control parameter of the angle of attack holding loop.

[0137] (2.3) Design of power compensation circuit based on constant speed

[0138] Because the angle of attack of the aircraft is kept constant under the control of direct lift, and the speed and track angle are decoupled, the throttle has the ability to maintain a constant speed. Therefore, the throttle is chosen to keep the speed constant.

[0139]

[0140] Among them, δ T0 is the throttle trim value, V is the actual speed, V c is the speed reference command, K p,V is the proportional control parameter of the power compensation loop, K i,V is the integral control parameter of the power compensation loop, K d,V is the differential control parameter of the power compensation loop.

[0141] (4) Conducting a glide path control simulation based on anti-disturbance control

[0142] Active disturbance rejection control is a control method that does not rely on the object model. The tracking differentiator (TD) is used to extract the continuous signal and differential signal of the system input; the extended state observer (ESO) is used to estimate the system state and total disturbance; the state error feedback (SEF) is used to design the control law according to the output of the extended state observer to achieve accurate control of the system state. The present invention introduces the attitude loop based on constant angle of attack and the power compensation loop based on constant speed into the active disturbance rejection control, which are described in detail below.

[0143] (4.1) Attitude loop based on constant angle of attack

[0144] In the direct lift control mode, the angle of attack needs to be kept constant, and the angle of attack holding circuit controls the angle of attack by introducing the elevon auxiliary control surface.

[0145] According to the aircraft dynamics formula, the following relationship is obtained:

[0146]

[0147] Writing the derivative relationship for the angle of attack:

[0148]

[0149] in, is the first-order derivative of the pitch angle, is the first derivative of the angle of attack.

[0150] Linearizing it at the reference point into the following equation:

[0151]

[0152] Substitute it into the dynamic model and we get:

[0153]

[0154] After sorting, we can get the following relationship:

[0155]

[0156] Among them, in order to comprehensively consider the first-order inertial link transfer function of the elevator aileron actuator, it is expanded and rewritten as:

[0157]

[0158] in, is the second-order derivative of the angle of attack, f 1 and b 1 Represents a relationship for the middle term.

[0159] Expanding the total disturbance in the angle of attack holding loop into a new state variable, the system is further expressed as:

[0160]

[0161] Among them, α 1 , α 2 , α 3 To expand variables, is the first-order derivative of the corresponding expanded variable, f 1 The first derivative of .

[0162] The extended state observer designed for the above system is as follows:

[0163]

[0164] Where: It is for α 1 An accurate estimate of Yes An accurate estimate of is an estimate of the total disturbance; β11 , β 12 and β 13 is the extended state observer gain of the angle of attack holding loop, β 11 =3ω 1 , ω 1 For bandwidth.

[0165] The following control law is obtained by introducing the tracking differentiator and state error feedback:

[0166]

[0167] Where: e α1 is the angle of attack error signal, e α2 is the angle of attack differential error signal, u 1 is the error feedback control output, α c is the desired tracking signal at angle of attack, is the desired differential signal of the angle of attack; is the proportional control parameter of the angle of attack holding loop, is the differential control parameter of the angle of attack holding loop.

[0168] Since the deflection of the trailing edge flap will produce some pitching moment, it is necessary to introduce the deflection of the elevon for compensation and decoupling, as shown in the following formula:

[0169]

[0170] Where: K is the torque decoupling coefficient.

[0171] (4.2) Power compensation circuit based on constant speed

[0172] The velocity equation is linearized as follows:

[0173]

[0174] in, is the first-order derivative of velocity.

[0175] The aircraft's thrust engine model is related to the throttle:

[0176]

[0177] Substitute it into the dynamic model and we get:

[0178]

[0179] After rearranging, we can get the following formula:

[0180]

[0181] Since the transfer function of the throttle lever actuator is also a first-order inertia link, the throttle lever transfer function is rewritten by comprehensively considering it:

[0182]

[0183] in, is the second-order derivative of velocity, f 2 and b 2 Represents a relationship for the middle term.

[0184] Consider the total disturbance in the dynamic compensation loop as an expanded state variable:

[0185]

[0186] Among them, V 1 、V 2 、V 3 To expand variables, is the first-order derivative of the corresponding expanded variable, f 2 The first derivative of .

[0187] The third-order active disturbance rejection control law is established as follows:

[0188]

[0189] Where: is the speed error signal, is the speed differential error signal, is the speed estimation error signal, V c and is the desired tracking and differentiation signal of the velocity; Yes V 1 estimates; yes estimates; is an estimate of the total disturbance; β 21 , β 22 and β 23 is the extended state observer gain of the dynamic compensation loop; β 21 =3ω 2 , ω 2 is the bandwidth; δ T is the throttle command; is the proportional control parameter of the power compensation loop, It is the differential control parameter of the power compensation loop.

[0190] (5) Conduct comparative simulation analysis

[0191] In order to verify the effectiveness of the direct lift anti-disturbance control method for flying wing layout proposed in this invention, comparative simulation experiments were carried out under the same simulation conditions. First, the direct lift anti-disturbance control method was compared and analyzed without interference, and then the angle of attack interference factor was introduced to conduct comparative experiments to verify the performance of this method in complex environments.

[0192] (5.1) Comparative simulation of active anti-disturbance glide control based on direct lift

[0193] Comparative simulation of the anti-disturbance glide control based on direct lift, such as Figure 6a , 6b , 6c, 6d, 6e, 6f, 6g, 6h.

[0194] By comparing the simulation experiments, it can be seen that the proposed method does not change significantly in terms of altitude deviation compared with the PID controller, but the changes in indicators such as track angle, pitch angle, and angle of attack are smoother. From the simulation results, the new design method performs better in terms of speed response and attitude stability. Specifically, the speed overshoot of the proposed method is reduced by 70.5%, the track angle overshoot is reduced by 28.2%, and the amplitude of change under the influence of disturbance is significantly lower than that of other control methods. The simulation results show that the designed anti-disturbance control method has significant advantages in terms of speed response and attitude stability, and can effectively suppress the influence of external disturbances on flight control.

[0195] (5.2) Comparative simulation of anti-disturbance glide path control under angle of attack disturbance

[0196] The comparative simulation in step (5.1) illustrates the advantages of ADRC. The following is to further verify the interference compensation of ADRC. The interference comparison simulation analysis of the two controllers, ADRC and PID, is performed. The interference introduction is shown as follows:

[0197] α disturb =0.5sint

[0198] Among them, α disturb is the introduced angle of attack interference, and t is the time.

[0199] Comparative simulation of active anti-disturbance glide control under angle of attack disturbance Figure 7a , 7b , 7c, 7d.

[0200] Through comparative simulation experiments, it can be seen that after adding ADRC control to the speed loop and angle of attack loop, the system state expansion observer effectively observes and compensates for the total disturbance of the system. The speed overshoot is reduced by 70%, and the angle of attack overshoot is reduced by 30%. The experimental results show that the height error is 0.56m and the landing point error is 0.41m, which meets the control accuracy requirements.

Claims

1. A direct lift active disturbance rejection control method suitable for a flying wing configuration, characterized in that: The details are as follows: (1) Dynamics modeling of flying wing aircraft During the aircraft's descent, the dynamic model of the aircraft's basic longitudinal motion is: Where: V is speed, T is thrust, α is angle of attack, D is drag, m is mass, g is acceleration due to gravity, γ is track angle, L is lift, I y is the moment of inertia, q is the pitch angular velocity, M is the pitch moment, x g is the forward displacement distance, H is the flight altitude, and θ is the pitch angle; Where: D is the drag, L is the lift, M is the pitching moment, C D is the drag coefficient, C L is the lift coefficient, C m is the pitching moment coefficient, ρ is the atmospheric density, S is the wing area, and c is the mean aerodynamic chord length; Where: PF is the trailing edge flap, δ ELE For the elevons, is the zero deviatoric drag derivative, is the drag derivative of the angle of attack, is the elevon drag derivative, is the drag derivative of the trailing edge flap, is the zero deflection lift derivative, is the lift derivative of the angle of attack, is the lift derivative of the elevon, is the lift derivative of the trailing edge flap, C m0 is the zero-bias pitch moment derivative, C mα is the pitching moment derivative of the angle of attack, is the trailing edge flap pitch moment derivative, is the elevon pitching moment derivative, H is the height; (2) Descent control design based on direct lift The PID control law is used in the altitude loop and the direct lift control loop. The control laws of the altitude loop and the direct lift control loop are as follows: The trailing edge flap is selected as the direct lift control surface to control the track angle; in the process of solving the control command, the track angle rate command can be calculated according to the deviation of the descent process, and the track angle rate can be directly controlled by controlling the change of the trailing edge flap to eliminate the deviation of the track angle; Among them, δ PF0 is the trailing edge flap trim value, is the actual value of the track angular velocity, is the reference track angular rate, K p,H is the proportional control parameter of the trajectory loop controller, K i,H is the integral control parameter of the trajectory loop controller, K d,H is the differential control parameter of the trajectory loop controller, is the proportional control parameter of the direct lift loop controller, is the integral control parameter of the direct lift loop controller, H is the actual height, H c is the height reference instruction, s is the complex variable of Laplace transform; (3) Design of Active Disturbance Rejection Control Method The anti-disturbance control method is adopted for the attitude loop based on constant angle of attack and the power compensation loop based on constant speed, as follows: (3.1) Attitude loop based on constant angle of attack In direct lift control mode, the angle of attack needs to be kept constant, and the angle of attack holding circuit controls the angle of attack by introducing the auxiliary control surface of the elevon; According to the aircraft dynamics formula, the following relationship is obtained: The derivative relationship of the angle of attack is as follows: in, is the first-order derivative of the pitch angle, is the first derivative of the angle of attack; Linearizing the angle of attack at the reference point into the following equation: Substitute it into the dynamic model and we get: The following relationship is obtained: Among them, in order to comprehensively consider the first-order inertial link transfer function of the elevator aileron actuator, it is expanded and rewritten as: in, is the second-order derivative of the angle of attack, f1 and b1 are intermediate terms that represent the relationship; Expanding the total disturbance in the angle of attack holding loop into a new state variable, the system is further expressed as: Among them, α1, α2, and α3 are extended variables. is the first-order derivative of the corresponding expanded variable, is the first-order derivative of f1; The extended state observer designed for the above system is as follows: Where: is an accurate estimate of α1; Yes An accurate estimate of is an estimate of the total disturbance; β 11 , β 12 and β 13 is the extended state observer gain of the angle of attack holding loop, β 11 =3ω1, ω1 is the bandwidth; The following control law is obtained by introducing the tracking differentiator and state error feedback: Where: e α1 is the angle of attack error signal, e α2 is the angle of attack differential error signal, u1 is the error feedback control output, α c is the desired tracking signal at angle of attack, is the desired differential signal of the angle of attack; is the proportional control parameter of the angle of attack holding loop, is the differential control parameter of the angle of attack holding loop; Since the deflection of the trailing edge flap will produce some pitching moment, it is necessary to introduce the deflection of the elevon for compensation and decoupling, as shown in the following formula: Where: K is the torque decoupling coefficient; (3.2) Power compensation circuit based on constant speed The velocity equation is linearized as follows: in, is the first-order derivative of velocity; The aircraft's thrust engine model is related to the throttle: Substitute it into the dynamic model and we get: After rearranging, we can get the following formula: Since the transfer function of the throttle lever actuator is also a first-order inertia link, the throttle lever transfer function is rewritten by comprehensively considering it: in, is the second-order derivative of velocity, f2 and b2 are intermediate terms that represent the relationship; Consider the total disturbance in the dynamic compensation loop as an expanded state variable: Among them, V1, V2, and V3 are extended variables. is the first-order derivative of the corresponding expanded variable, is the first-order derivative of f2; The third-order active disturbance rejection control law is established as follows: Where: is the speed error signal, is the speed differential error signal, is the speed estimation error signal, V c and is the desired tracking and differentiation signal of the velocity; is the estimate of V1; yes estimates; is an estimate of the total disturbance; β 21 , β 22 and β 23 is the extended state observer gain of the dynamic compensation loop; β 21 =3ω2, ω2 is the bandwidth; δ T is the throttle command; is the proportional control parameter of the power compensation loop, It is the differential control parameter of the power compensation loop.

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