A tilt-rotor aircraft transition state control strategy based on incremental dynamic inverse

Abstract

The application discloses a tilt-rotor aircraft transition state control strategy based on incremental dynamic inverse. On the basis of establishing a nonlinear system model of the tilt-rotor aircraft, a new control strategy is provided, that is, a PID control law is used in an outer loop speed circuit and acceleration instruction distribution is used, and an incremental dynamic inverse (INDI) control law is used in an inner loop attitude angle circuit. The INDI method does not depend on an accurate aircraft dynamics model, and has excellent control ability of other adaptive nonlinear control methods. The control strategy has superiority for the tilt-rotor aircraft to smoothly and safely transit from a helicopter mode to a fixed-wing mode under the condition of a fixed height. The control system designed in the application can well meet the demand of transition control, and provides an effective solution for the tilt-rotor aircraft transition state control.

Description

Technical Field

[0001] This invention belongs to the field of flight control, and more specifically relates to a transitional control strategy for tiltrotor aircraft based on incremental dynamic inversion. Background Technology

[0002] Tiltrotor aircraft, as complex aircraft capable of vertical takeoff and landing and high-speed horizontal flight, have long been a focus of research in the aviation field. Due to their combination of helicopter and fixed-wing aircraft characteristics, they require switching between different control modes, which undoubtedly increases the complexity of control. Especially during transitions—that is, switching from helicopter mode to fixed-wing mode, or vice versa—extremely high demands are placed on the design and implementation of control strategies.

[0003] On the other hand, incremental dynamic inverse control is a control strategy that has been widely used in control theory in recent years. Its main advantage is that it provides a unified framework for controller design and analysis, making the design and analysis of control systems simpler and more universal. Secondly, this control method does not require obtaining an accurate nonlinear model of the controlled object, such as a tiltrotor aircraft. However, its application in the transient flight control problem of tiltrotor aircraft is limited, and many problems still need to be studied and solved.

[0004] Current research on transition states mainly focuses on two aspects: first, the study of transition state theory, primarily delving into the definition, property analysis, and methods for better understanding and describing transition states; second, the study of control strategies under transition states, mainly addressing control problems in transition states and designing stable and efficient control strategies. However, due to the highly complex dynamic model of tiltrotor aircraft, and the significant changes in the system's dynamic characteristics caused by switching control modes during transition states, the study of control strategies under transition states has always been a very challenging problem. Summary of the Invention

[0005] This invention proposes a PID + acceleration command allocation + INDI control scheme, employing a PID control law and acceleration command allocation in the outer loop velocity loop and an incremental dynamic inverse (INDI) control law in the inner loop attitude angle loop. This scheme enables transition control of the time-varying and complex control process of a tiltrotor aircraft transitioning from helicopter mode to fixed-wing mode. The INDI method does not rely on a precise aircraft dynamics model, yet it possesses the excellent control capabilities of other adaptive nonlinear control methods.

[0006] To achieve the above objectives, the present invention employs the following technical solution: the control strategy includes:

[0007] A multibody dynamics model of a tiltrotor aircraft is established, the multibody motion characteristics of the tiltrotor aircraft are analyzed, a nonlinear model of the tiltrotor aircraft is established, and a hybrid control model of the tiltrotor aircraft is established.

[0008] Control plan and command generation for the transition phase of a tiltrotor aircraft, where only longitudinal maneuvers occur, including:

[0009] Establish the engine nacelle tilt law;

[0010] Establish the target state function of the tiltrotor aircraft. During the transition from helicopter mode to airplane mode, the tiltrotor aircraft needs to maintain its flight altitude and complete the flight mode transition in a monotonically increasing manner within the flight envelope.

[0011] The design incorporates a flight control law for the transition phase of a tiltrotor aircraft, employing a multi-loop control structure. It establishes an attitude controller and acceleration command distribution module based on incremental dynamic inversion to control the tiltrotor aircraft to achieve a safe and smooth transition from helicopter mode to fixed-wing mode under constant altitude conditions.

[0012] Furthermore, the nonlinear model of the tiltrotor aircraft is established using a Lagrange-based multi-rigid-body modeling method.

[0013] Furthermore, the aforementioned rotorcraft hybrid control model has two control modes: helicopter control mode and airplane control mode.

[0014] In helicopter mode, the tiltrotor aircraft achieves flight control through collective control pitch, longitudinal cyclic pitch, lateral cyclic pitch, and differential pitch.

[0015] In aircraft mode, tiltrotor aircraft achieve flight control by manipulating engine throttle, ailerons, elevators, and rudders;

[0016] Both control methods are effective in helicopter mode, but the aircraft control methods are not obvious at low speeds. In transition mode, the engine nacelle gradually tilts from helicopter mode to aircraft mode, and the helicopter control methods gradually stop. In aircraft mode, the tiltrotor aircraft relies entirely on the aircraft control methods.

[0017] Furthermore, the aforementioned engine nacelle tilt law adopts a tilt mode with an engine nacelle tilt angle of "uniform acceleration - uniform speed - uniform deceleration";

[0018] The tilting law of the engine nacelle of a tiltrotor includes two tilting modes, which correspond to: (1) transitioning from helicopter mode to airplane mode; (2) transitioning from airplane mode to helicopter mode.

[0019] Furthermore, the target state function of the tiltrotor is a function of the flight speed with respect to the nacelle tilt angle; in order to maintain altitude during mode transition, the expected value of the vertical velocity should be constant at zero throughout the tilt transition process.

[0020] Furthermore, in the aforementioned multi-loop control structure, the outer loop uses a PID control law and acceleration command allocation to generate the desired pitch angle and collective pitch angle; the inner loop uses an incremental dynamic inverse INDI control law to track the desired attitude.

[0021] Furthermore, the incremental dynamic inverse INDI control law does not directly calculate the total control input signal, but instead calculates the increment of the control input based on the given system conditions within each sampling time.

[0022] Furthermore, the outer loop employs a PID control law and acceleration command allocation;

[0023] The horizontal acceleration command is obtained through the PID controller based on the tracking error of the horizontal forward velocity; the vertical acceleration command is obtained through the PID controller based on the tracking error of the vertical velocity.

[0024] In helicopter mode, the tiltrotor's forward acceleration is provided by the pitch angle, and the vertical acceleration is provided by the rotor.

[0025] In fixed-wing mode, the forward acceleration of the tiltrotor aircraft is provided by the rotor, and the vertical acceleration is provided by the pitch angle;

[0026] The transitional flight phase is mixed.

[0027] Beneficial effects of this invention:

[0028] This invention proposes a transitional control strategy for tiltrotor aircraft based on incremental dynamic inversion. This method does not rely on a precise aircraft dynamics model and exhibits superior robustness to system uncertainties and external disturbances, ensuring the stability of the tiltrotor's flight altitude throughout the transition process and significantly reducing the pilot's workload. This control strategy enables the tiltrotor to smoothly and safely transition from helicopter mode to fixed-wing aircraft mode while maintaining stable flight altitude. This research provides a novel and practically feasible solution for the design of control strategies for tiltrotor aircraft, possessing significant theoretical and practical value. Attached Figure Description

[0029] Figure 1 These are the three flight modes for tiltrotor aircraft;

[0030] Figure 2 This is a schematic diagram of the longitudinal dynamics model of a tiltrotor aircraft.

[0031] Figure 3 Two control modes for tiltrotor aircraft;

[0032] Figure 4 It is a combination of two control methods;

[0033] Figure 5 The tilting pattern of the engine nacelle;

[0034] Figure 6 The flight envelope and three tilt transition curves;

[0035] Figure 7 This is a structural diagram of a multi-loop control system;

[0036] Figure 8 For speed tracking curves;

[0037] Figure 9 This is the attitude angle tracking curve;

[0038] Figure 10 This is the curve showing the change in the manipulated quantity. Detailed Implementation

[0039] To make the objectives, technical solutions, and beneficial effects of the present invention clearer, the preferred embodiments of the present invention will be described in detail below to facilitate understanding by those skilled in the art.

[0040] The control strategy includes:

[0041] A multibody dynamics model of a tiltrotor aircraft was established to analyze the multibody motion characteristics of the tiltrotor aircraft.

[0042] A tiltrotor aircraft is a variable-structure aircraft that combines the characteristics of fixed-wing aircraft and helicopters. Unlike conventional fixed-wing aircraft and helicopters, a tiltrotor aircraft is a complex multi-body system. Depending on the angle of its nacelle, it can operate in three modes: helicopter mode, transition mode, and aircraft mode. Figure 1 As shown.

[0043] By tilting the engine nacelles, tiltrotor aircraft can switch between helicopter and airplane flight modes, combining the advantages of both. In helicopter mode, the tiltrotor can fly freely in different directions and achieve vertical takeoff and landing as well as hovering. In airplane mode, if the tiltrotor's flight speed is high enough that the lift generated by the wings is sufficient to overcome the aircraft's gravity, the rotor can act as a propeller for a fixed-wing aircraft, providing forward propulsion and enabling high-speed flight.

[0044] Establish a nonlinear model of the tiltrotor aircraft

[0045] like Figure 2As shown, a longitudinal dynamic model of the tiltrotor aircraft is established in the ground coordinate system. Point O is the position of the rotation axis of the engine nacelle-rotor system, point A is the position of the rotor hub, OA is the axis of the engine nacelle and is perpendicular to the hub plane; G1 represents the position of the fuselage center of mass, G2 represents the position of the engine nacelle-rotor system center of mass, and G represents the position of the overall aircraft center of mass; δ AF Let δ be the angle between straight line OG1 and the longitudinal axis of the aircraft fuselage. TM θ is the angle between straight line OG2 and the engine nacelle axis; θ is the pitch angle of the tiltrotor aircraft. N L is the tilt angle of the engine nacelle, V is the airspeed of the tiltrotor aircraft, and α is the angle of attack of the tiltrotor aircraft; L, D, and M m These represent the lift, drag, and pitching moment experienced by the wing-body assembly, respectively. α 1s L is the rear chamfer of the rotor tip plane relative to the rotor hub plane. T The thrust is from the rotor. In addition, the tiltrotor aircraft is also subject to the gravity of the fuselage and engine nacelle-rotor system, as well as the control torque M that rotates the engine nacelle-rotor system. N The positive direction of this torque is the same as the positive direction of the pitching torque.

[0046] In addition, the rotor will also generate a tilt angle b under the action of the incoming flow. 1s When only considering the longitudinal dynamics of the tiltrotor aircraft, the two rotors are identical in all parameters except for their direction of rotation. Therefore, due to the different sideslip angles β, the two rotors will have different rotational speeds. 1s The forces and torques generated can cancel each other out.

[0047] Based on the Lagrange equations, select the position (x, z) of the midpoint O of the rotation axis in the ground coordinate system, the fuselage pitch angle θ, and the nacelle tilt angle θ. N As a generalized coordinate, let θ AF =θ-δ AF θ TM =θ+θ N -δ TM The expression for generalized force can be obtained as follows:

[0048] Q1=Lsin(α-θ)-Dcos(α-θ)+L T cos(α 1s +θ+θ N )

[0049] Q2=Lcos(α-θ)+Dsin(α-θ)+L T sin(α 1s +θ+θ N )-mg

[0050] Q3 = M m +m1l1(Lsinα-Dcosα+mgcosθAF ) / m

[0051] +m2l2[Lcos(α+θ N )+Dsin(α+θ N )-mgcosθ TM ] / m

[0052] +l3L T sina 1s

[0053] Q4 = M N +m2l2[Lcos(α+θ N )+Dsin(α+θ N )-mgcosθ TM ] / m+l3L T sina 1s (1)

[0055] Wherein, the generalized coordinates corresponding to Q1, Q2, Q3 and Q4 are x, z, θ and θ, respectively. N l1, l2, and l3 represent the lengths of OG1, OG2, and OA, respectively; m1, m2, and m represent the mass of the fuselage, engine nacelle-rotor system, and the entire aircraft, respectively.

[0056] Substituting the tilt law of the engine nacelle into the longitudinal multibody dynamics model of the tiltrotor aircraft can eliminate the control variable M. N .make By adding a kinematic equation, the longitudinal dynamic model of the tiltrotor aircraft can be written in the following form.

[0057]

[0058] Where q represents the pitch angular velocity;

[0059]

[0060] a 13 =a 31 =m1l1cosθ AF -m2l2sinθ TM

[0061] a 23 =a 32 =m1l1sinθ AF +m2l2cosθ TM

[0062] a 33 =I y1 +I y2 +m1l1 2+m2l2 2

[0063]

[0064] Among them, I y1 and I y2 These are the moments of inertia of the fuselage and the engine nacelle-rotor system, respectively.

[0065] Based on the aerodynamic theory of blades, calculate L T and α 1s as follows:

[0066]

[0067] Where Ω,θ0 and B1 represent the rotor speed, root collective pitch and longitudinal cyclic pitch, respectively.

[0068] The aerodynamic coefficients are derived from wind tunnel test data from Bell Labs and NASA's Langley Laboratory. Furthermore, helicopters and transition-mode wings are affected by rotor wake. Therefore, neglecting rotor contribution, the overall aerodynamic forces and moments can be expressed as follows:

[0069]

[0070] Where V i δ and k represent the induced velocity and influence coefficient of the rotor, respectively; e It refers to the elevator angle.

[0071] It is worth noting that in B1 and δ e There is a control coupling between them. They are both controlled by the displacement δ of the longitudinal control rod. long control.

[0072] Therefore, we can obtain four state variables v that describe the longitudinal plane motion of the tiltrotor aircraft. x v z q and θ, and two control variables θ0 and δ long Let X = [v] x ,v z ,q,θ] T U = [θ0, δ long ] T Then equation (2) can be written in the general form of a nonlinear system of equations.

[0073]

[0074] Establish a hybrid control model for rotorcraft

[0075] Tiltrotor aircraft have two control modes: helicopter control and airplane control. In helicopter mode, the tiltrotor achieves flight control through collective pitch, longitudinal cyclic pitch, lateral cyclic pitch, and differential pitch. In airplane mode, the tiltrotor achieves flight control through engine throttle, ailerons, elevators, and rudder. Both control modes are effective in helicopter mode, but the airplane control mode is less noticeable at low speeds. In transition mode, as the nacelle gradually tilts from helicopter mode to airplane mode, the helicopter control mode gradually ceases. In airplane mode, the tiltrotor relies entirely on airplane control. The two control modes of the tiltrotor are as follows: Figure 3 As shown.

[0076] For the pilot, the control mechanisms are the collective pitch throttle, longitudinal control stick, lateral control stick, and pedals. The control values ​​for helicopter control and aircraft control should correspond to their respective channels. In helicopter control, vertical movement is controlled by the rotor collective pitch, corresponding to the collective pitch throttle; pitch movement is controlled by the rotor longitudinal cyclic pitch, corresponding to the longitudinal control stick; roll movement is controlled by the rotor differential collective pitch, corresponding to the lateral control stick; and yaw movement is controlled by the rotor differential longitudinal cyclic pitch, corresponding to the pedals. In aircraft control, rotor thrust is controlled by the throttle, corresponding to the collective pitch throttle; pitch movement is controlled by the elevator, corresponding to the longitudinal control stick; roll movement is controlled by the ailerons, corresponding to the lateral control stick; and yaw movement is controlled by the rudder, corresponding to the pedals. The specific correspondence between the two control methods and the control channels is as follows: Figure 4 As shown.

[0077] Tiltrotor Transition Section Control Plan and Command Generation

[0078] The aerodynamic complexity of tiltrotor aircraft in transition mode is mainly reflected in their longitudinal dynamic characteristics. Furthermore, to ensure the stability of the tiltrotor aircraft, it typically only performs longitudinal maneuvers during transition flight. Therefore, this invention focuses on the design of a longitudinal motion controller. The lateral control of the tiltrotor aircraft is consistent with that of conventional aircraft, employing a traditional inner-loop control method.

[0079] Establish engine nacelle tilt law

[0080] Properly tilting the engine nacelle not only allows the tiltrotor to quickly switch flight modes, but also prevents the tiltrotor from losing control due to excessively rapid changes in aerodynamic forces.

[0081] The multibody dynamics model of the tiltrotor aircraft derived in this invention involves the angular acceleration of the nacelle tilt angle. This paper describes the nacelle tilt angle as a tilt mode of "uniform acceleration - uniform speed - uniform deceleration".

[0082]

[0083] As shown in equation (6), the nacelle tilt angle is π / 2 in helicopter mode and 0 in airplane mode, where a and b are constants. The tilt law of the engine nacelle of the tiltrotor aircraft includes two tilt modes, corresponding to: 1. transition from helicopter mode to airplane mode; 2. transition from airplane mode to helicopter mode. Through multiple simulation experiments, this invention finally selected the values ​​of a and b under the two tilt modes, as shown in Table 1.

[0084] Table 1 Selection of Engine Nacelle Tilting Law Parameters

[0085]

[0086] Table 1 shows that when the engine nacelle tilts in the first manner, the entire tilting process takes 18 seconds; when the engine nacelle tilts in the second manner, the entire tilting process takes 33 seconds. The nacelle tilting pattern during the transition from helicopter mode to fixed-wing mode in a tiltrotor aircraft is shown in [Table 1]. Figure 5 .

[0087] Establish the target state function for the tiltrotor aircraft. During the transition from helicopter mode to airplane mode, the tiltrotor aircraft needs to maintain its flight altitude and complete the flight mode transition in a monotonically increasing manner within its flight envelope.

[0088] The target state refers to the desired motion state of the tiltrotor aircraft, which can also be considered as the controller reference command. The purpose of designing the controller is to make the actual state of the tiltrotor aircraft as close as possible to the target state. When the tiltrotor aircraft transitions from helicopter mode to airplane mode, the desired flight state during the mode transition is to maintain altitude and complete the flight mode transition within the flight envelope in a monotonically increasing velocity manner. Below, we analyze the flight envelope and generate the target state within it.

[0089] The flight envelope describes the safe flight speed range of a tiltrotor aircraft at different nacelle tilt angles. The flight envelope of the XV-15 is as follows: Figure 5 As shown in the figure, the horizontal axis represents the flight speed of the tiltrotor aircraft, the vertical axis represents the nacelle tilt angle, and the boundaries are obtained by calculating the maximum and minimum flight speeds of the tiltrotor aircraft in stable forward flight at a certain fixed nacelle tilt angle.

[0090] A tiltrotor aircraft transitions from helicopter mode to airplane mode via numerous possible transition tilt curves within its flight envelope. For example, if the flight speed in helicopter mode is V0 = 30 m / s, and the flight speed in airplane mode is V... T =90m / s, Figure 6 Three tilt transition curves are given in the figure, and their numbers are shown in the figure.

[0091] like Figure 6 As shown, each transition tilt curve can be expressed as the flight speed V with respect to the nacelle tilt angle θ. N The function. Meanwhile, equation (6) describes the nacelle inclination angle θ. N A function of time t. From this, we can obtain the target state function of flight speed V with respect to time t, i.e.

[0092] V Ref =V(θ) N )=V(θ N (t)) (7)

[0093] Taking the NO.1 tilt transition curve as an example, the specific expression of the target state function of flight speed is as follows:

[0094]

[0095] The velocity in the longitudinal plane can be decomposed into v x and v z Two components: vertical velocity v, which is used to maintain altitude during mode transitions in the tiltrotor. z The expected value should remain zero throughout the entire tilt transition process, i.e.

[0096] v z,Ref (t)≡0 (9)

[0097] v x,Ref (t)=V Ref (t) (10)

[0098] The design incorporates a flight control law for the transition phase of the tiltrotor aircraft, employing a multi-loop control structure to ensure a safe and smooth transition from helicopter mode to fixed-wing mode at a constant altitude.

[0099] exist Figure 7 In the middle, v x v z v x,Ref and v z,Ref The definitions are the horizontal velocity, vertical velocity, desired horizontal velocity, and desired vertical velocity of the tiltrotor aircraft, respectively. x,Ref and a z,Ref These are the expected horizontal and vertical accelerations of the tiltrotor aircraft during the transition phase, respectively. θ Ref It is the desired pitch angle generated by the acceleration command allocation module, φ Ref and ψ Ref These are the desired roll angle and yaw angle, respectively, which are set to 0. θ 0,RefIt is the desired rotor root collective pitch. U is defined as the same as U in equation (5) and is generated by the INDI attitude controller. U0 is the control quantity given to the tiltrotor after the actual action of the actuator.

[0100] Establish an attitude controller based on incremental dynamic inverse

[0101] INDI does not directly calculate the total control input signal, but rather calculates the increment of the control input based on the given system conditions within each sampling time interval. At time Δt, equation (5) can be rewritten as a first-order Taylor series expansion:

[0102]

[0103] Where ΔX = X - X0 and ΔU = U - U0 represent the increments of X and U within a sampling time step Δt, respectively; X0 and U0 represent X and U at the previous sampling time, respectively; Δ H.O.T This indicates a higher-order term. It can be expressed by the following formula:

[0104]

[0105] When the control update frequency is high enough (time increment Δt is extremely small), U will change significantly faster than X, and the value of X will approach X0 while the value of U will not approach U0. Therefore, equation (11) can be simplified to:

[0106]

[0107] Replace with pseudo-inverse control input v Rearranging equation (13) yields the following control law for INDI:

[0108]

[0109] Substituting equation (14) into equation (13) yields

[0110]

[0111] Therefore, the control signal U in equation (14) linearizes the nonlinear system equation (13), and the linearized control law is:

[0112] Establish an acceleration command allocation module

[0113] According to v x The tracking error is obtained by the PID controller to obtain the forward acceleration command a. x,Ref According to v z The tracking error is obtained by the PID controller through the vertical acceleration command a. z,RefIn helicopter mode, the tiltrotor's forward acceleration is provided by the pitch angle, and its vertical acceleration is provided by the rotor; in fixed-wing mode, the tiltrotor's forward acceleration is provided by the rotor, and its vertical acceleration is provided by the pitch angle; however, in the transition phase, it is a mixture of both. Therefore, this invention uses a sine function allocation method to generate the pitch angle command θ. Ref And the rotor's expected collective root pitch θ 0,Ref :

[0114]

[0115] Where, θ N It is the engine nacelle tilt angle, which changes according to the engine nacelle tilt rotation law.

[0116] To address the complex time-varying process of a tiltrotor aircraft transitioning from helicopter mode to fixed-wing mode, simulations were performed using the transition control plan and transition flight control law designed in this invention, ultimately yielding the following results: Figure 8 , Figure 9 and Figure 10 The results are shown. Figure 8 and Figure 9 The results show that, under the control of the controller designed in this invention, the tiltrotor aircraft achieved tracking of the desired speed and desired pitch angle. Specifically, the mean and variance of the horizontal speed tracking error are 0.0553 m / s and 1.9650 m / s, respectively. 2 / s 2 The mean and variance of the vertical velocity tracking error are 0.0197 m / s and 0.1712 m, respectively. 2 / s 2 The mean and variance of the pitch tracking error are 0.0040 degrees and 0.2067 degrees, respectively. 2 The mean and variance of these tracking errors are very small, indicating that under the transition control plan and transition flight control law designed in this invention, the tiltrotor aircraft smoothly transitioned from helicopter mode to fixed-wing mode. Figure 10 As can be seen, the control quantities used during the transition process are all within a reasonable range and meet the actual usage requirements.

[0117] Those skilled in the art will understand that all or part of the processes in the above embodiments can be implemented by a computer program instructing related hardware. The program can be stored in a computer-readable storage medium, and when executed, it can include the processes of the embodiments of the above methods. The storage medium can be a magnetic disk, optical disk, read-only memory (ROM), or random access memory (RAM), etc.

[0118] It should be understood that the above detailed description of the technical solutions of the present invention with reference to preferred embodiments is illustrative and not restrictive. Those skilled in the art can modify the technical solutions described in the embodiments or make equivalent substitutions for some of the technical features based on reading this specification; however, these modifications or substitutions do not cause the essence of the corresponding technical solutions to depart from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A transitional control strategy for tiltrotor aircraft based on incremental dynamic inversion, characterized in that: The control strategy includes: A multibody dynamics model of a tiltrotor aircraft is established, the multibody motion characteristics of the tiltrotor aircraft are analyzed, a nonlinear model of the tiltrotor aircraft is established, and a hybrid control model of the tiltrotor aircraft is established. Control plan and command generation for the transition phase of a tiltrotor aircraft, where only longitudinal maneuvers occur, including: Establish the engine nacelle tilt law; Establish the target state function of the tiltrotor aircraft. During the transition from helicopter mode to airplane mode, the tiltrotor aircraft needs to maintain its flight altitude and complete the flight mode transition in a monotonically increasing manner within the flight envelope. The design of the transition flight control law for the tiltrotor aircraft adopts a multi-loop control structure and establishes an attitude controller and acceleration command distribution module based on incremental dynamic inverse to control the tiltrotor aircraft to achieve a safe and smooth transition from helicopter mode to fixed-wing mode under constant altitude conditions. The aforementioned engine nacelle tilt law adopts a tilt mode with an engine nacelle tilt angle of "uniform acceleration - uniform speed - uniform deceleration"; The tilt law of the engine nacelle of the tiltrotor includes two tilting modes, which correspond to: (1) transitioning from helicopter mode to airplane mode; (2) transitioning from airplane mode to helicopter mode. The multi-loop control structure described above uses a PID control law and acceleration command allocation in the outer loop to generate the desired pitch angle and collective pitch angle; the inner loop uses an incremental dynamic inverse INDI control law to track the desired attitude. The incremental dynamic inverse INDI control law does not directly calculate the total control input signal, but calculates the increment of the control input based on the given system conditions in each sampling time. The outer loop employs PID control law and acceleration command allocation; The horizontal acceleration command is obtained through a PID controller based on the tracking error of the horizontal forward velocity; the vertical acceleration command is obtained through a PID controller based on the tracking error of the vertical velocity. In helicopter mode, the tiltrotor's forward acceleration is provided by the pitch angle, and the vertical acceleration is provided by the rotor. In fixed-wing mode, the forward acceleration of the tiltrotor aircraft is provided by the rotor, and the vertical acceleration is provided by the pitch angle; The transitional flight phase is mixed.

2. The tiltrotor transient control strategy based on incremental dynamic inversion according to claim 1, characterized in that: The nonlinear model of the tiltrotor aircraft was established using a Lagrange-based multi-rigid-body modeling method.

3. The tiltrotor transient control strategy based on incremental dynamic inversion as described in claim 1, characterized in that: The aforementioned rotorcraft hybrid control model has two control modes: helicopter control mode and airplane control mode. In helicopter mode, the tiltrotor aircraft achieves flight control through collective control pitch, longitudinal cyclic pitch, lateral cyclic pitch, and differential pitch. In aircraft mode, tiltrotor aircraft achieve flight control by manipulating engine throttle, ailerons, elevators, and rudders; In transition mode, the engine nacelle gradually tilts from helicopter mode to airplane mode, and helicopter control gradually ceases. In aircraft mode, tiltrotor aircraft rely entirely on aircraft control methods.

4. The tiltrotor aircraft transient control strategy based on incremental dynamic inversion according to claim 1, characterized in that: The target state function of the tiltrotor is the flight speed as a function of the nacelle tilt angle; in order to maintain altitude during mode transition, the expected value of the vertical velocity should be zero throughout the tilt transition process.

Images