Tilting dual-rotor aircraft attitude control method

Abstract

The invention discloses a tilt dual-rotor aircraft attitude control method, which belongs to the technical field of unmanned aerial vehicle control, and comprises the following steps: establishing a dynamic model with center-of-gravity offset and a propeller disc coordinate system; a reference coordinate system fed back by attitude control of the tilting dual-rotor aircraft is transferred from an aircraft body coordinate system to a propeller disc coordinate system; the angle control ring adopts a paddle disc inclination angle of a paddle disc surface relative to the ground as negative feedback; the angular velocity control loop adopts the body angular velocity under a paddle disc coordinate system as angular velocity negative feedback; finally, the angular velocity control loop is subjected to proportional differentiation and then is subjected to feed linear superposition with a paddle disc inclination angle to be output to a dynamical model of a motor and a steering engine, and a control loop is achieved. According to the tilting dual-rotor aircraft, the minimum phase problem and the gravity center offset instability problem of an existing tilting dual-rotor aircraft are solved, the gravity center offset resisting capacity is improved, robustness is high, and response is fast.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of tilt-rotor aircraft, in particular to a tilt-rotor aircraft control method with anti-center-of-gravity shift capability and minimum phase characteristic. BACKGROUND

[0002] The tilt-rotor aircraft is an aircraft that changes the thrust direction of the propeller / rotor plane by rotating the propeller / rotor, thereby achieving position / attitude control. This type of aircraft changes the type of aircraft actuator, has strong accessibility, fast flight speed, long endurance time, vertical take-off and landing, and flat flight, and has broad application prospects in military and civilian fields.

[0003] However, there are two control problems in the current system: the tilt-rotor aircraft with the center of gravity below the propeller plane is proved to be a non-minimum phase system in the attitude loop, and the use of conventional control methods based on pitch feedback will produce adverse phenomena of reverse overshoot; secondly, the tilt-rotor aircraft is difficult to cope with external disturbances, such as changes in the center of gravity caused by load movement. These challenges limit its further promotion and application.

[0004] The existing solution to the non-minimum phase problem mainly solves it in the aspects of machinery and control, but there are still shortcomings. Mechanically, the propeller plane can be designed to be below the center of gravity, which can make the aircraft angle loop become a minimum phase system, but this method cannot improve the non-minimum phase phenomenon of the angular velocity loop, and the inverted motor will also make the rotor disc closer to the ground, and the disturbance of near-ground turbulence will be significantly intensified, which will worsen the control stability. Another mechanical optimization method is to change the propeller structure and use periodic variable pitch passive hinge control, without introducing rudder to adjust the thrust direction. This innovation eliminates the nonlinearity, lag and non-minimum phase problems caused by the rudder, however, this rudder-free dual-rotor aircraft model will produce vibration due to mechanical inaccuracy in practical application, which affects the measurement accuracy of IMU, and may cause resonance, facing the difficulty of overall promotion.

[0005] The following will analyze the non-minimum phase problem of the tilt-rotor aircraft using the existing control method.

[0006] The tilt-rotor aircraft, as shown in Figure 1 , includes a rotor, a motor, a rudder and a rocker arm, a carbon fiber rod, a base, and a battery. The thrust of the tilt-rotor aircraft is provided by the motor driving the rotor, and the total thrust is T, and the thrust of the left and right rotors is T1 and T2. The rudder is connected to the motor, and by rotating the rudder, the motor can be driven to achieve the purpose of tilting the rotor. According to the platform, let denote the body coordinate system fixed with the center of gravity, denote the world coordinate system, both of which adopt the north-east-ground right-hand coordinate system. The attitude between the body coordinate system and the world coordinate system is represented by a rotation matrix Defined as follows, where Euler angles Θ = [φ, θ, ψ] T Roll, pitch, and yaw are represented in the ZXY order. The angular velocity of the tiltrotor aircraft is described as ω = [p, q, r]. T The two motors are directly driven by servos mounted at the ends of the V-frame, ensuring that the rotor tilt angle equals the servo tilt angle. The rotational speeds of the left and right motors are Ω1 and Ω2, respectively. The thrust of the left and right propellers and the disk tilt angle (servo deflection angle) are represented as T1, δ1, T2, and δ2, respectively, parallel to the XOZ coordinate system. D represents the distance between the servo rotation axis and the x-axis, and L is half the platform width.

[0007] The dynamic modeling in the body coordinate system is as follows. The equations describe the force and torque balance relationship during aerial motion, forming the basis for the motion analysis of the tiltrotor aircraft:

[0008]

[0009] Where: m and J B Let G represent the mass and moment of inertia of the aircraft in body coordinates, respectively; p represents the position of the center of gravity relative to the world coordinate system; and G = [0, 0, mg] is the moment of inertia along the body coordinate system. W Gravity along the z-axis J is the rotation matrix between the coordinate system and the world coordinate system. yy To circle y B Moment of inertia of the shaft, B ω is the angular velocity in the body coordinate system, τ B and B T represents the external torque and total thrust in the body coordinate system, respectively, and is calculated using the following equations (3) and (4):

[0010] τ B =τ T +τ m,s +τ p1 +τ p2 +τ g1 +τ g2 (3);

[0011] Where: τ T τ represents the torque generated by the rotor thrust. m,s τ represents the counter-torque of the motor-servo actuator assembly. p1 With τ p2 τ represents the rotor torque generated by the rotation of two propellers. g1 With τ g2 This represents the gyroscopic torque generated by the change in spatial position when the two propellers rotate.

[0012] Similar to most tiltrotor aircraft designs, the total thrust generated by the rotor thrust is...B T and torque τ T Represented as:

[0013]

[0014] Where D represents the servo motor rotation axis and y B The distance between the axes, L is half the width of the platform, and T1, δ1 and T2, δ2 represent the thrust of the two propellers and the deflection angle of the propeller disk, respectively.

[0015] The rotor torque τ in equation (3) p1 ≈K Ω Ω1 2 ,τ p2 ≈-K Ω Ω2 2 The torque generated by the thrust vector is negligible due to the counter-rotating propellers. To ensure modeling accuracy, the gyroscopic torque generated by the high-speed rotation of the motor and servo motor is considered.

[0016] The following are basic constraints on the flight conditions. The tiltrotor aircraft performs attitude control under the given flight altitude h = h0, pitch angle θ = θ(t), yaw angle φ = 0, and roll angle ψ = 0. Considering that the yaw and roll angles are preset to 0, substituting into formula (2) yields the independent pitch motion along y B The moment balance equation for the shaft is:

[0017]

[0018] (6) J yy To circle y B Moment of inertia of the shaft, To circle y B Angular acceleration of the axis, τ T [2,1] is the term in the second row and first column of the moment matrix in formula (5). (7) In formula J m,s The moment of inertia of the motor-servo actuator assembly. This is the angular acceleration of the servo motor's deflection.

[0019] Substituting the result of equation (7) into equation (6) and performing a Laplace transform on equation (6), we obtain the following expression, which serves as the theoretical basis for system attitude control:

[0020] J yy s 2 θ=mgDδ-J m,s s 2 (δ+θ) (8);

[0021] Regarding the non-minimum phase problem of the attitude loop in dual-rotor UAVs, for a conventional dual-rotor UAV system (such as dual-rotor A), if a conventional pitch attitude angle is used as the control feedback, the following equation applies:

[0022]

[0023] Where δ des For the desired servo deflection angle, θ des For the desired fuselage pitch angle, θ, K represents the actual pitch angle, pitch velocity, and pitch acceleration of the fuselage. pθ K pq K dq These are adjustable parameters that can be dynamically adjusted based on the actual system.

[0024] And assume that the input and output relationship of the servo motor satisfies the following second-order system:

[0025]

[0026] Where ζ and ω n These are the damping ratio and natural frequency of the second-order system, respectively.

[0027] Combining equations (8), (9), and (10), and using a small-angle linear approximation with sinθ≈θ and cosθ≈1, we obtain the final input and output transfer functions:

[0028]

[0029] Where den is the denominator polynomial of the transfer function, K pθ K pq K dq It is an adjustable parameter consistent with equation (9), where s is the complex frequency and is the independent variable in the complex frequency domain analysis.

[0030] Setting the numerator polynomial of the transfer function to 0, since the parameter is actually used, K... pq With K pθ The value is not 0, and the value of the natural frequency is also not 0, only mgD-s 2The value of the expression may be 0. Therefore, when finding the zeros of the transfer function, a positive zero will appear in the right half-plane. Studies have shown that the appearance of such a zero in the s-plane leads to the system exhibiting non-minimum phase characteristics. Specifically, given a desired pitch angle, a tiltrotor aircraft will experience forward oscillation around the servo axis due to inertia during forward flight. This phenomenon misleads the flight control system into misjudging the flight direction, thus applying excessive correction commands to the rotor, causing it to deviate from the preset state. The control system then generates new correction values, forming a continuous positive feedback loop, ultimately causing the fuselage to oscillate continuously in the air and disrupting the stability of the roll velocity loop. This is the fundamental cause of the non-minimum phase problem.

[0031] Some studies have attempted to mitigate or avoid the impact of the aforementioned control problems by modifying control strategies. A common solution is to introduce a highly robust control loop into the controller, enhancing the stability margin of the tiltrotor aircraft through weighted function optimization. However, the controller's order is too high, resulting in a slow system response. Even after reducing the order and sacrificing control stability, the problem remains unresolved. In actual flight verification, the dual rotors require more than 20 seconds to converge, far below the agility expected of a dual rotor system. This suboptimal control performance is primarily attributed to the control method's insufficient consideration of the unique dynamic characteristics of the tiltrotor platform.

[0032] The second challenge of tiltrotor platforms stems from their center-of-gravity (CG) offset instability. Due to their unique configuration, twin-rotor aircraft must dynamically adjust their attitude during disturbances to ensure the thrust line passes through the CG. This change in steady-state attitude angle is related to an uncertain CG offset. This characteristic makes conventional controller methods for combating CG offset—such as mechanical structure optimization, adaptive control algorithms, cascaded control, and robust design—unsuitable for direct application to twin-rotors. Furthermore, the CG offset in the pitch direction of a tiltrotor cannot be solved solely by differential motor control; it requires servo angle adjustments, but existing control methods cannot quickly and accurately calculate these adjustments. Most research on twin-rotor aircraft fixes the CG in a specific position—this simplification leads to suboptimal control performance in actual twin-rotor aircraft.

[0033] These problems reduce the expected endurance and flexibility of the dual-rotor system, making a new modeling and control method a goal pursued by those skilled in the art. Summary of the Invention

[0034] The purpose of this invention is to address the common attitude loop minimum phase problem and center of gravity offset instability problem in tilt-rotor aircraft, and to design a highly versatile attitude control method for tilt-rotor aircraft that eliminates the non-minimum phase problem of the attitude loop, improves its resistance to center of gravity offset, and achieves a combination of high stability, high agility and long endurance.

[0035] To achieve the above-mentioned objectives, the technical solution adopted by this invention is as follows: a tiltrotor aircraft attitude control method, characterized in that the control method comprises: establishing a dynamic model with a center of gravity offset and a rotor disk coordinate system; transferring the reference coordinate system for attitude control feedback of the tiltrotor aircraft from the body coordinate system to the rotor disk coordinate system; using the rotor disk tilt angle relative to the ground as negative feedback in the angle control loop; using the body angular velocity in the rotor disk coordinate system as negative feedback in the angular velocity control loop; and finally, the angular velocity control loop, after proportional-derivative analysis, is linearly superimposed with the rotor disk tilt angle feedforward and output to the dynamic model of the motor and servo motor to realize the control loop.

[0036] In the tilt-rotor aircraft attitude control method of the present invention, as a preferred embodiment, the angle control loop uses the rotor disk tilt angle as negative feedback, that is, θ+δ as negative feedback. Since the servo motor rotation angle cannot be directly obtained, i.e., the rotor disk tilt angle cannot be known, the present invention adds a servo motor rotation angle observation module, i.e., a servo motor rotation angle observer, which obtains the estimated rotor disk tilt angle by fitting through a built-in PID observer. As the propeller disk tilt angle, the control error of the angle control loop at this time is:

[0037] Among them, e θ For angle control error, θ is the pitch angle, θ des The expected value of the pitch angle. This is the estimated value of the propeller disk tilt angle.

[0038] The tilt-rotor aircraft attitude control method of the present invention, as a preferred embodiment, wherein the estimated value of the rotor disk tilt angle... It is obtained through the following calculation method:

[0039] The input disk tilt angle expectation value δ of the servo motor angle observation module des After passing through the proportional control module in the servo angle observation module, the expected value of the propeller disk tilt angle δ is determined. des Error e of propeller disk tilt angle δ The relationship is as follows:

[0040]

[0041] Among them, K P δ is an adjustable parameter of the proportional control module. des e is the desired value of the propeller disk tilt angle. δ This is the propeller disk tilt angle error;

[0042] The torque equation is calculated using a proportional-integral control model:

[0043]

[0044] Among them, K I K P These are the adjustable parameters of the proportional and integral modules, respectively, where τ is the torque; and the propeller disk tilt angle error e δ and the error of its first derivative can be and Find the answer.

[0045] The torque equation, calculated using the principles of dynamics, is as follows:

[0046]

[0047] Among them: J m,s Let τ be the moment of inertia of the motor-servo actuator assembly, and τ be the torque. The estimated value of the propeller disk tilt angle involved in formula (21) can be obtained from the following formula (22):

[0048]

[0049] Combining formulas (20) and (21), we obtain the desired value of the propeller disk tilt angle δ. des The expression for the satisfied differential equation is:

[0050]

[0051] K in the formula P ,K I The values ​​can be obtained by fitting experiments based on different servo motor and motor combinations. A Hall element with a rotating shaft can be used to measure the real-time servo motor rotation angle. When the fitting degree is close, the angle measured by the servo motor rotation angle observation module can be considered acceptable. The propeller disk tilt angle is the estimated value that can be calculated using equation (23).

[0052] The tilt-rotor aircraft attitude control method of the present invention, as a preferred embodiment, wherein the desired value of the rotor disk tilt angle δ des It is obtained by superimposing the expected pitch angle value input in the pitch direction through the proportional control module, the proportional-derivative control module, and the feedforward proportional amplification module, that is, by calculating the following formula:

[0053]

[0054] Where, q des K is the angular velocity in the pitch direction. pq K dq K ff These are the adjustable parameters θ and θa of the proportional control module, the proportional derivative module, and the feedforward proportional amplification module, respectively. des δ represents the actual pitch angle and the expected pitch angle of the fuselage.des Let the desired pitch angle of the propeller disk be denoted as . This is the pitch acceleration.

[0055] The tilt-rotor aircraft attitude control method of the present invention, as a preferred embodiment, wherein the angular velocity loop control method is as follows:

[0056] First, based on the pitch angle control error, the expected value of the pitch direction angular velocity is obtained through proportional control, as shown in the following formula:

[0057] q des =K pθ e θ (twenty three);

[0058] Where: q des K represents the expected value of the angular velocity in the pitch direction. pθ These are the adjustable parameters of the proportional control module;

[0059] Secondly, the body angular velocity observed in the propeller disk coordinate system is used as control feedback, and the body angular velocity ω in the body coordinate system is multiplied by the rotation matrix on the left. BW Transformed to the propeller disk coordinate system, the body angular velocity ω disk The calculation formula is:

[0060]

[0061] The negative feedback error of the body angular velocity is:

[0062] e q =q des -ω disk [2,1] (25);

[0063] Where: e q The angular velocity error in the pitch direction, ω is the estimated value of the propeller disk tilt angle. body ω is the angular velocity of the machine body in the machine system. disk Let ω be the angular velocity of the airframe under the propeller disk system. disk [2,1] The term in the second row and first column of the angular velocity matrix in the propeller disk coordinate system.

[0064] The attitude control method for tilt-rotor aircraft described in this invention addresses the adverse effects of center of gravity offset on the stability of a tilt-rotor UAV. Therefore, this invention considers the influence of center of gravity offset when establishing the dynamic model. Under ideal conditions, a new dynamic model is established. In this new model, the center of gravity, mass, and moment of inertia of the entire aircraft are affected by the external weight in the pitch direction, C′,m′,J. yyThe parameters introduced to simulate the center of gravity offset represent the new center of gravity position, the mass of the external load, and the moment of inertia of the new model about the y-axis, respectively. The new dynamic model is:

[0065]

[0066] J yy ′ is the moment of inertia of the aircraft about the y-axis after the center of gravity is offset, m is the mass of the aircraft, m′ is the mass of the external load, l is the distance of the external load, g is the acceleration due to gravity, T is the total thrust, D represents the distance between the servo motor rotation axis and the x-axis, δ is the propeller disk tilt angle, θ is the pitch angle, and when the external load is attached, α is the deflection angle of the external pylon relative to the horizontal attachment.

[0067] The tiltrotor aircraft attitude control method described in this invention, as a preferred embodiment, assumes a servo motor.

[0068] A second-order system whose input and output relationships satisfy the following:

[0069]

[0070] Where ζ and ω n Here, δ represents the damping ratio and natural frequency of the second-order system, respectively; s is the independent variable in the complex frequency domain analysis, representing the complex frequency; δ is the propeller disk tilt angle, δ des Let G(s) be the expected value of the propeller disk tilt angle. The transfer function between the expected value and the actual value of the propeller disk tilt angle is denoted as G(s).

[0071] Combining formulas (15), (27), and (28), and using a small-angle linear approximation of sinθ≈θ and cosθ≈1, the transfer function between the control input and output can be obtained:

[0072]

[0073] in:

[0074] num=(K ff +K pθ K pq )[(J yy ′-J s )s 2 +LT]ω n 2 (30);

[0075]

[0076] Where: num is the numerator transfer function polynomial, den is the denominator transfer function polynomial; θ is the pitch angle, δ is the propeller disk tilt angle, δ des K represents the desired pitch angle of the propeller disk. ff K pθ Kpq K dq J is an adjustable parameter for the feedforward proportional amplifier module, proportional control module, and proportional derivative module. yy ′ is the moment of inertia about the y-axis, J s Let be the moment of inertia of the servo motor, s be the complex frequency (the independent variable in the complex frequency domain analysis), L be half the platform width, T be the total thrust, and ζ and ω be the independent variables. n These are the damping ratio and natural frequency of the second-order system, respectively.

[0077] Setting the numerator transfer function polynomial of formula (30) to 0, it is easy to see that the real part of the resulting zero is 0, located on the imaginary axis. It can be seen that after introducing the propeller tilt angle control strategy, the positive zero point located in the right half plane no longer appears in the transfer function of this invention, proving that the control method described in this invention has solved the problem of the non-minimum phase zero point, thereby significantly improving the stability of the entire control system.

[0078] This invention significantly improves control stability by transforming the collective coordinate system into a propeller disk coordinate system and selecting the propeller disk tilt angle as negative feedback, effectively solving the minimum phase problem of the attitude loop and the instability problem caused by center of gravity offset. This is because this invention is inspired by the output redefinition and zero-point tuning optimization method, and optimizes stability by constructing control feedback under a virtual reference frame. In this model, the virtual coordinate system is the propeller disk plane coordinate system. This invention redefines the input relationship, transforming the traditional relationship of input fuselage pitch angle - output desired fuselage pitch angle under the airframe system into the sum of input fuselage pitch angle and propeller disk tilt angle - output desired fuselage pitch angle under the propeller disk plane coordinate system. After applying this definition, the original non-minimum phase zero point will be transferred to the imaginary axis of the complex plane, eliminating the influence of the original system's non-minimum phase zero point, thereby improving the system's stability. At the same time, based on the analysis of the airframe's steady-state characteristics and physical derivation, we note that regardless of the center of gravity offset, the quadrotor aircraft can achieve a steady-state pitch angle of zero on the airframe platform. When extending this characteristic to tilt-rotor UAVs, we found that when the center of gravity shifts uncertainly, the pitch angle of the aircraft in steady state, i.e., the angle measured by the flight control gyroscope, will change uncertainly. However, the tilt angle of the propeller disk relative to the ground is always zero in steady state. Therefore, this invention utilizes this key feature to transform the reference frame of the dual-rotor UAV to the propeller disk plane coordinate system, that is, to allow the flight control gyroscope to measure the angle change in the propeller disk system, so that the tilt-rotor aircraft can achieve the same good control stability effect as the quadcopter aircraft. Attached Figure Description

[0079] Figure 1 This is a schematic diagram of the tilt-rotor aircraft described in this invention;

[0080] Figure 2This is a force analysis diagram of the tilt-rotor aircraft described in this invention, showing the center of gravity shift caused by the addition of counterweight.

[0081] Figure 3 This is a block diagram of the attitude control of the tilt-rotor aircraft described in this invention;

[0082] Figure 4 This is a block diagram of the servo motor angle observation module in the attitude control of the tiltrotor aircraft described in this invention;

[0083] Figure 5 This is a comparison chart showing the effectiveness of the propeller tilt angle negative feedback method described in this invention and existing control methods in solving the non-minimum phase problem;

[0084] Figure 6 This is a comparison chart showing the control effects of the propeller tilt angle negative feedback method described in this invention and existing control methods in solving the center of gravity offset problem. Detailed Implementation

[0085] The preferred embodiments of the present invention will be described in detail below with reference to the accompanying drawings, but these preferred embodiments should not be used to limit the scope of protection of the present invention.

[0086] The tilt-rotor aircraft used in this invention, in addition to simulating the external load and related parameters of the center of gravity offset, also includes the settings of the rotor, motor, servo motor and rocker arm, carbon fiber rod, base, and battery. Figure 1 As shown, it also includes: rotors, motors, servos and rocker arms, carbon fiber rods, a base, and batteries. The thrust of the tilt-rotor aircraft is provided by the motors driving the rotors, with a total thrust of T. The thrust of the left and right rotors is T1 and T2, respectively. The servos are connected to the motors; rotating the servos drives the motors to achieve the tilting of the rotors. Based on this platform, let... This represents the body coordinate system fixed to the center of gravity. The world coordinate system is represented, and both are North-East-Ground right-handed coordinate systems. The attitude between the body coordinate system and the world coordinate system is determined by a rotation matrix. Defined as follows, where Euler angles Θ = [φ, θ, ψ] T Roll, pitch, and yaw are represented in the ZXY order. The angular velocity of the tiltrotor aircraft is described as ω = [p, q, r]. T The two motors are directly driven by servos mounted at the ends of the V-frame, ensuring that the rotor tilt angle equals the servo tilt angle. The rotational speeds of the left and right motors are Ω1 and Ω2, respectively. The thrust of the left and right propellers and the disk tilt angle (servo deflection angle) are represented as T1, δ1, T2, and δ2, respectively, parallel to the XOZ coordinate system. D represents the servo rotation axis relative to the x-axis. B The distance between axes, L is half the width of the platform.

[0087] To address the issues of center of gravity shift and non-minimum phase, this invention employs the following control method: A dynamic model is established under ideal conditions with added counterweight and center of gravity offset; the reference coordinate system for control feedback of the tiltrotor aircraft is transferred from the body coordinate system to the propeller disk coordinate system; the angle control loop uses the propeller disk tilt angle relative to the ground as negative feedback; the angular velocity control loop uses the body angular velocity in the propeller disk coordinate system as negative angular velocity feedback; finally, the angular velocity control loop, after proportional-derivative processing, is linearly superimposed with the propeller disk tilt angle feedforward and output to the motors and servos to realize the control loop.

[0088] First, the shift in the center of gravity should be considered when establishing the dynamic model; see [reference needed]. Figure 2 Add external loads to the fuselage, where T1+T2 is the rotor thrust, δ is the rotor disk tilt angle, θ is the pitch angle, C′ is the center of gravity position after adding external loads, m′ is the mass of the external loads, m is the fuselage mass, C is the original center of gravity, and x B Let z be the X-axis of the body coordinate system. B is the Z-axis of the body coordinate system.

[0089] Considering the offset of the center of gravity, the dynamic equations under ideal conditions are as follows:

[0090]

[0091] J yy Let m' be the moment of inertia of the new model considering the weight around the y-axis, m' be the mass of the externally mounted weight, l be the distance of the externally mounted weight, g be the acceleration due to gravity, D be the distance between the servo motor's rotation axis and the x-axis, L be half the platform width, δ be the propeller disk tilt angle, θ be the pitch angle, and α be the deflection angle of the external mounting rod relative to the horizontal mounting when the external weight is set. Specifically, when the weight of the externally mounted weight is small, the external mounting rod can be considered to have no deformation, i.e., the weight is exactly horizontal. In this embodiment, α is assumed to be 0 degrees, simplifying the subsequent modeling and analysis process.

[0092] Under the condition θ≈0, by using small-angle approximations of sinθ≈θ and cosθ≈1, we can obtain the linearized differential equation:

[0093]

[0094] The quantities represented by the coefficients K1, K2, and K3 are as follows:

[0095] K1 = J yy ′+m′l 2 ,K2=TD,K3=-m′glcosα (17);

[0096] Among them: J yy ′ is the moment of inertia about the y-axis;

[0097] Solving the differential equation (16) determines the new equilibrium state vector as follows:

[0098]

[0099] From the result of equation (18), it can be intuitively seen that when the aircraft finally reaches steady-state hovering, the fuselage has a non-zero pitch angle, and this angle is related to the initial setting conditions of the center of gravity offset. Since the initial conditions of the center of gravity offset are uncertain in engineering practice, i.e., parameters such as m′ and l are unknown, the value of the fuselage pitch angle at the final stable state cannot be obtained from this result. However, at the level of control theory, the above control equation reflects an important relationship—regardless of whether the center of gravity offset exists or not, the conclusion δ+θ=0 always holds true in steady-state hovering.

[0100] When the system is in steady-state hovering, the equation δ + θ = 0 is essentially a mathematical representation of the physical principle that the thrust line passes through the center of gravity. That is, when the dual-rotor UAV reaches steady state, its rotor disk plane always remains horizontal to the ground.

[0101] Based on this, the tiltrotor aircraft attitude control method of the present invention designs an angle control loop and an angular velocity control loop negative feedback loop, and simultaneously, an improved control method based on propeller disk tilt angle (servo motor rotation angle) feedforward. Specifically, the present invention transfers the control feedback reference coordinate system of the tiltrotor aircraft from the body coordinate system (or flight control fixed system) to the propeller disk coordinate system, and performs negative feedback control of the angle loop and angular velocity loop. The angle control loop uses the propeller disk tilt angle as negative feedback, and the angular velocity control loop uses the body angular velocity in the propeller disk coordinate system as angular velocity negative feedback. In addition, feedforward control based on angle error is added in parallel to the servo motor input stage, and finally output to the motor and servo motor to realize the control loop.

[0102] See Figure 3 The diagram below shows the control block diagram of the tilt-rotor aircraft attitude control method of the present invention. The controller can be divided into the following main parts: an angle negative feedback loop based on the rotor disk tilt angle, an angular velocity loop negative feedback loop based on the linear superposition of servo angle feedforward and angular velocity loop, a servo angle observation module, and a fuselage / servo dynamics part.

[0103] See Figure 3 First, the angle control loop of the present invention: when the desired pitch angle θ is input... des Then, the propeller tilt angle δ is used as negative feedback. At this point, the control error is... Instead of e θ =θ des -θ, this is one of the key principles of the present invention, which uses the relative tilt angle between the propeller disk and the ground, i.e., the propeller disk tilt angle δ, as negative feedback.

[0104] For the measurement of θ+δ, since the propeller tilt angle δ, i.e., the servo motor angle, cannot be directly obtained, therefore, see... Figure 3 , 4 This invention adds a servo motor angle observation module for predicting the servo motor angle, i.e., the propeller disk tilt angle δ. Figure 4 This is the control block diagram of the servo motor angle observation module; specifically, the estimated propeller tilt angle is obtained by using a built-in PID observer for fitting, combined with the rotational inertia, lever arm and parameters of the servo motor and motor.

[0105] Input the desired pitch angle θ des After passing through the proportional control module, the desired value δ of the propeller disk tilt angle is obtained. des The relationship between the pitch angle error and the propeller disk tilt angle error is as follows:

[0106]

[0107] Where: K P e is an adjustable parameter of the proportional control module in the servo motor angle observation module. δ δ represents the pitch angle error of the propeller disk. des Let be the expected value of the propeller disk tilt angle, and let be the estimated value of the propeller disk tilt angle. Equation (19) represents the relationship between the expected value of the propeller disk tilt angle and the propeller disk tilt angle error after passing through the proportional control module. This equation can establish the connection between the expected value of the propeller disk tilt angle and the actual value.

[0108] The torque can be calculated using a proportional-integral control model (i.e., based on the error quantity of feedback control), and the solution formula is as follows:

[0109]

[0110] Among them, K I K P These are the coefficients of the proportional and integral modules, respectively, and τ is the torque. The propeller disk tilt error e... δ and the error of its first derivative and can be and Find the answer.

[0111] The torque is calculated using the principles of dynamics as follows:

[0112]

[0113] J m,s Let τ be the moment of inertia of the motor-servo actuator assembly, and τ be the torque. The estimated value of the propeller disk tilt angle involved in formula (21) can be obtained from the following formula (22):

[0114]

[0115] Combining formulas (20) and (21), the differential equation expression satisfying the desired value of the propeller disk tilt angle can be obtained as follows:

[0116]

[0117] For K in equation (20) P ,K I The value can be obtained by fitting experiments based on different servo and motor combinations. A Hall element with a rotating shaft can be used to measure the real-time propeller tilt angle. When the degree of fitting is close to the target value, the angle measured by the servo angle observer can be considered as the target value. It refers to the actual propeller disk tilt angle δ.

[0118] After obtaining the attitude control error (i.e., the control error of the angle control loop). Then, the angular error is proportionally controlled to obtain the expected value of the angular velocity:

[0119] q des =K pθ e θ (twenty four);

[0120] Where q des K represents the angular velocity in the pitch direction. pθ These are the adjustable parameters of the proportional control module.

[0121] In the angular velocity loop, this invention uses the body angular velocity observed in the propeller disk coordinate system as the control feedback, as shown in equation (25) below. ω is multiplied by the rotation matrix on the left. body Converting to the angular velocity in the propeller disk coordinate system yields the negative feedback of the angular velocity:

[0122]

[0123] At this point, the angular velocity error is:

[0124] e q =q des -ω disk [2,1] (26);

[0125] In the formula, e q The angular velocity error in the pitch direction, ω is an estimated value for the propeller disk tilt angle. body Let ω be the angular velocity in the body coordinate system. disk Let ω be the angular velocity in the propeller disk coordinate system. disk [2,1] represents the term in the second row and first column of the angular velocity matrix in the propeller disk coordinate system.

[0126] angular velocity error e q After proportional-integral control and initial angle deviation e θThe desired tilt angle δ of the propeller disk is obtained by feedforward superposition. des The output is then fed to the motor and servo motor via a linear superposition converter. Feedforward control is used because the desired steering angle of the servo motor is actually related to e. θ Relatedly, without adding feedforward, the controller may develop near-steady-state errors due to the increased complexity of the angle-angular velocity control chain, which is undesirable in actual flight. Therefore, the initial angle deviation e θ Feedforward control solves the problem of the potential for steady-state errors caused by the complexity of the angle-angular velocity control chain.

[0127] Specifically, the desired propeller disk tilt angle δ can be calculated from the input fuselage pitch angle in the control loop. des See the following formula:

[0128]

[0129] Equation (27) reflects how the desired value of the fuselage pitch angle θ is determined in the control loop. des The desired value δ of the propeller disk tilt angle was calculated. des In other words, it establishes the relationship between the propeller disk tilt angle and the fuselage pitch angle under the desired conditions. In this formula, q des Let θ be the expected value of the angular velocity in the pitch direction, and θ be the expected value of the angular velocity in the des For the actual pitch angle and the expected pitch angle of the fuselage, δ and δ des K represents the actual tilt angle of the propeller disk and the expected tilt angle of the propeller disk; where K pq K dq With K ff These are the adjustable parameters for the proportional module, the derivative module, and the feedforward module, respectively. K pq K dq With K pθ K dθ The relationship is similar to that between q and θ, where the former is a quantity in the pitch direction of the latter.

[0130] And assume that the input and output relationship of the servo motor satisfies the following second-order system:

[0131]

[0132] Where ζ and ω n Here, δ represents the damping ratio and natural frequency of the second-order system, respectively; S is the independent variable in the complex frequency domain analysis, representing the complex frequency; and δ is the propeller disk tilt angle. des Let G(s) be the desired propeller tilt angle. For simplicity and convenience, we denote the transfer function between the desired and actual servo angle as G(s).

[0133] Combining equations (15), (27), and (28), and using a small-angle linear approximation of sinθ≈θ and cosθ≈1, and considering body dynamics, we can obtain the transfer function between the expected input and output of the pitch angle considering the center of gravity offset:

[0134]

[0135] in:

[0136] num=(K ff +K pθ K pq )[(J yy ′-J s )s 2 +LT]ω n 2 (30);

[0137]

[0138] Where num is the numerator transfer function polynomial and den is the denominator transfer function polynomial. Similar to equation (11), setting the numerator transfer function polynomial of equation (29) to 0, it is easy to see that the real part of the resulting zero is 0, located on the imaginary axis. It can be seen that after introducing the paddle disk tilt angle as a negative feedback control strategy in this invention, the positive zero point located in the right half-plane no longer appears in its transfer function, proving that the control method of this invention has solved the non-minimum phase problem and has a very considerable improvement on the stability of the entire control system.

[0139] See Figure 4 The figure compares the performance of the propeller tilt angle negative feedback control described in this invention with existing control methods in solving the non-minimum phase problem. The figure shows a comparison of the control performance and input of the traditional PID control method and the propeller tilt angle negative feedback control method of this invention. It is clearly evident from the figure that the control method of this invention has a faster response speed and shorter settling time, and does not exhibit reverse overshoot beyond the input, effectively solving the attitude loop non-minimum phase problem, demonstrating strong robustness and rapid response.

[0140] Also see Figure 5The diagram compares the control performance of the propeller tilt angle negative feedback control method described in this invention with existing control methods in solving the center of gravity offset problem. The blue, green, and red lines in the diagram represent the traditional PID control method, the ACA control method, and the propeller tilt angle negative feedback control method of this invention, respectively. The experiment was set as hover control, with a center of gravity offset step input provided at one second. The response time, steady-state time, overshoot, and final steady-state effect of the aircraft under the three control methods were evaluated. A comprehensive comparison shows that the traditional PID control method not only lacks an advantage in reaching steady-state time, but also that both the propeller tilt angle and fuselage pitch angle are 0 when steady-state is reached. This means that the aircraft is actually maintaining a constant acceleration in forward flight in steady-state. In other words, the traditional PID control method fails to achieve a return to a stable hovering state, i.e., the PID control method lacks the ability to counteract center of gravity offset. The ACA control method and the propeller tilt angle negative feedback control method of this invention ultimately return the UAV to a stable hovering state, but the ACA method exhibits significant overshoot and oscillation phenomena, and requires 3 seconds to reach a stable state. The control method of the present invention responds quickly, reaching a stable state in about 0.5 seconds, and basically eliminates the reverse overshoot phenomenon, thus exhibiting excellent control performance.

[0141] Ultimately, the input motor speed is determined by directly adding or subtracting the roll angle control value from the hover throttle on both sides. Here, since the roll angle is zero, the input motor speed is directly the hover speed. The PWM values ​​(ESC duty cycles) input to the two servos are:

[0142]

[0143] PWM stands for Electronic Switch Duty Cycle, with a value ranging from 0 to 1, used to characterize the rotation angle of a servo motor. In equation (32), the ESC duty cycles corresponding to the two servos are... and It is only related to the desired servo deflection angle δ des Related quantities.

[0144] It should be noted that the attitude control method described in the above embodiments of the present invention is for controlling the pitch angle, but it is also applicable to controlling the yaw angle and roll angle.

[0145] The above description is illustrative only and not restrictive of the present invention. The present invention aims to provide an attitude control method for a tiltrotor aircraft, which transforms the body coordinate system to a propeller disk coordinate system for negative feedback control of angle loops and angular velocity loops. The angle loop uses the propeller disk tilt angle as negative feedback, and the angular velocity loop uses the angular velocity in the propeller disk coordinate system as negative feedback. Those skilled in the art will understand that many modifications, variations, or equivalents can be made without departing from the spirit and scope defined by the claims, and all such modifications, variations, or equivalents will fall within the protection scope of the present invention.

Claims

1. A method for attitude control of a tiltrotor aircraft, characterized in that, The control method is as follows: a dynamic model with a center of gravity offset and a propeller disk coordinate system are established; the reference coordinate system for attitude control feedback of the tiltrotor aircraft is transferred from the body coordinate system to the propeller disk coordinate system; the angle control loop uses the propeller disk tilt angle relative to the ground as negative feedback; the speed control loop uses the body angular velocity in the propeller disk coordinate system as negative feedback; finally, the angular velocity control loop is linearly superimposed with the propeller disk tilt angle feedforward after proportional-derivative analysis and output to the dynamic model of the motor and servo motor to realize the control loop.

2. The attitude control method for a tiltrotor aircraft according to claim 1, characterized in that, The angle control loop uses propeller tilt angle as negative feedback, specifically θ+δ, where the propeller tilt angle δ is an estimated value obtained by fitting the propeller tilt angle using the servo angle observation module. At this time, the control error of the angle control loop is: Among them, e θ For angle control error, θ is the pitch angle, θ des The expected value of the pitch angle. This is the estimated value of the propeller disk tilt angle.

3. The attitude control method for a tiltrotor aircraft according to claim 3, characterized in that, The estimated value of the propeller disk tilt angle Calculated using the following method: Input the desired pitch angle θ des After passing through the proportional control module, the desired value δ of the propeller disk tilt angle is obtained. des The relationship between the pitch angle error and the propeller disk tilt angle error is as follows: Among them, K P δ is an adjustable parameter of the proportional control module. des e is the desired value of the propeller disk tilt angle. δ This refers to the propeller disk tilt angle error. This is an estimated value for the propeller disk tilt angle; The torque equation is calculated using a proportional-integral control model: Among them, K I K P These are the adjustable parameters of the proportional-integral module, where τ is the torque; and the propeller disk tilt error e is the torque. δ The error of its first derivative can be and Find the answer. The torque equation, calculated using the principles of dynamics, is as follows: Among them: J m,s Let τ be the moment of inertia of the motor-servo actuator assembly, and τ be the torque. The estimated value of the propeller disk tilt angle involved in formula (21) can be obtained from the following formula (22): Combining formulas (20) and (21), we obtain the desired value of the propeller disk tilt angle δ. des The expression for the satisfied differential equation is: The estimated value of the propeller disk tilt angle can be calculated using this equation.

4. The attitude control method for a tiltrotor aircraft according to claim 3, characterized in that, The desired value of the propeller tilt angle δ des Calculated using the following formula: Where, q des K represents the expected value of the angular velocity in the pitch direction. pq K dq K ff These are the adjustable parameters θ and θa of the proportional control module, the proportional derivative module, and the feedforward proportional amplification module, respectively. des δ represents the actual pitch angle and the expected pitch angle of the fuselage. des This represents the desired pitch angle of the propeller disk.

5. The attitude control method for a tiltrotor aircraft according to any one of claims 1-4, characterized in that, The angular velocity loop control method is as follows: First, based on the pitch angle control error, the expected value of the pitch direction angular velocity is obtained through proportional control, as shown in the following formula: q des =K pθ e θ (23); Where: q des K represents the expected value of the angular velocity in the pitch direction. pθ These are the adjustable parameters of the proportional control module; Secondly, the body angular velocity observed in the propeller disk coordinate system is used as control feedback, and the body angular velocity ω in the body coordinate system is multiplied by the rotation matrix on the left. BW Transformed to the propeller disk coordinate system, the body angular velocity ω disk The calculation formula is: The negative feedback error of the body angular velocity is: e q =q des -oh disk [2,1] (25); Where: e q The angular velocity error in the pitch direction, ω is an estimated value for the propeller disk tilt angle. body Let ω be the angular velocity in the body coordinate system. disk Let ω be the angular velocity in the propeller disk coordinate system. disk [2,1] The item in the second row and first column of the angular velocity matrix in the propeller disk coordinate system is the angular velocity of the aircraft in the pitch direction as observed in the propeller disk coordinate system.

6. The attitude control method for a tiltrotor aircraft according to claim 1, characterized in that, After considering the center of gravity offset, the ideal dynamic model of the motor and servo motor is as follows: J yy ′ is the moment of inertia of the aircraft about the y-axis after the center of gravity is offset, m is the mass of the aircraft, m′ is the mass of the external load, l is the distance of the external load from the original center of gravity of the aircraft, g is the acceleration due to gravity, T is the total thrust, D represents the distance between the servo motor rotation axis and the x-axis, δ is the propeller disk tilt angle, θ is the pitch angle, and when the external load is attached, α is the angle formed by the line connecting the center of gravity of the load and the original center of gravity of the aircraft and the external pylon.

7. The attitude control method for a tiltrotor aircraft according to claim 6, characterized in that, Based on the empirical model, we assume that the input and output relationship of the servo motor satisfies the following second-order system: Where ζ and ω n Here, δ represents the damping ratio and natural frequency of the second-order system, respectively; s is the independent variable in the complex frequency domain analysis, representing the complex frequency; δ is the propeller disk tilt angle, δ des Let G(s) be the expected value of the propeller disk tilt angle. The transfer function between the expected value and the actual value of the propeller disk tilt angle is denoted as G(s). Combining equations (15), (27), and (28), and using a small-angle linear approximation of sinθ≈θ and cosθ≈1, the transfer function between the control input and output can be derived: in: num=(K ff +K pθ K pq )[(J yy ′-J s )s 2 +LT]ω n 2 (30); Where: θ is the pitch angle, δ is the propeller disk tilt angle, δ des K represents the desired pitch angle of the propeller disk. ff K pθ K pq K dq J is an adjustable parameter for the feedforward proportional amplifier module, proportional control module, and proportional derivative module. yy ′ is the moment of inertia about the y-axis, J s Let be the moment of inertia of the servo motor, s be the complex frequency (the independent variable in the complex frequency domain analysis), L be half the platform width, T be the total thrust, and ζ and ω be the independent variables. n These are the damping ratio and natural frequency of the second-order system, respectively.

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