Design method of longitudinal stabilization system in transition section of tiltrotor aircraft based on flying qualities

Abstract

The present invention provides a design method for a longitudinal stabilization system of a tiltrotor aircraft transition section based on flight quality, which belongs to the field of flight mechanics and flight simulation technology, and includes the following steps: S1: analyzing the longitudinal modal characteristics of the tiltrotor aircraft; S2: designing an ideal model based on flight quality requirements; designing an inner loop control law based on an incremental backstepping method; thereby constructing a longitudinal stabilization system; S3: evaluating the stabilization effect of the longitudinal stabilization system on the dynamic process of tilt transition. The present invention designs ideal models under different modes in combination with the requirements for the smoothness of the tilt transition. Based on the control concept of the incremental backstepping method, an inner loop attitude angular velocity control law with the pilot and the ideal model as the outer loop attitude angle is designed, which can be applicable to the stabilization requirements of manned tiltrotors in the tilt transition section. The tilt transition path is designed, and a pilot model is established to perform quality analysis to verify the effectiveness of the stabilization system design.

Description

Technical Field

[0001] The invention belongs to the technical field of flight mechanics and flight simulation, and in particular relates to a design method for a longitudinal stabilization system of a transition section of a tiltrotor aircraft based on flight quality. Background Art

[0002] Tilt-rotor aircraft are currently the most practical high-speed rotorcraft. By changing the inclination angle of the rotor nacelles at both ends of the wing, it can achieve the conversion between helicopter mode, fixed-wing aircraft mode, and tilt-rotor transition mode, thus combining the vertical take-off and landing, air hovering capabilities of helicopters and the high-speed cruising capabilities of fixed-wing aircraft. It has a wide range of application prospects. References [1,2]. However, switching flight modes brings great challenges to flight quality design.

[0003] The quality requirements of tilt-rotor aircraft in the three flight modes are different. The military helicopter flight quality specification ADS-33E-PRF reference [3] establishes standards for helicopter control stability, combining the helicopter's response form to control commands with stability to form response types such as RCDH, ACAH, and RCHH. Different levels of response types are used for flight missions of different difficulty and visual conditions to meet the first-level quality requirements. It is suitable for helicopter mode flight quality assessment. Military manned fixed-wing aircraft flight quality MIL-STD-1797A reference [4] constructs a first-level quality range based on longitudinal modal parameters and related indicators according to the longitudinal motion characteristics of fixed-wing aircraft, which is suitable for fixed-wing aircraft mode flight quality assessment. However, the flight quality specification for the time-varying system of the tilt-rotor transition section is not yet mature. The most commonly used one is the time-varying flight quality criterion constructed by Klyde based on the wavelet analysis method, which can be used for type III PIO prediction references [16,17,18].

[0004] In order to build a control system that adapts to the full-mode flight of a tiltrotor aircraft, it is necessary to establish an integrated tiltrotor dynamics model. Based on the aerodynamic modeling of each component, the aerodynamic interference between components, and the control redundancy analysis, Chen Renliang established a complete flight dynamics model of the XV-15

[13] . The model was verified by comparing the GTRS

[14] data, and the flight mode conversion method was studied to provide a basis for the selection of the tilt transition path.

[0005] In terms of control research, there have been many developments in recent years on automatic flight control. The LPV control system theory for variable structure aircraft with nonlinear time-varying characteristics converts complex nonlinear time-varying systems into linear parameter-dependent systems, so that advanced linear control methods can be applied to the controlled objects [5,6], reducing the difficulty of control system design. However, the parameters are highly dependent and quality switching is not possible. The nonlinear dynamic inverse control method is suitable for flight control strategy switching. Smooth switching is achieved by designing weight functions [7]. This method is applied to tilt-rotor quad-rotor UAVs, and control is achieved using a cascaded inner and outer loop control structure [8]. The incremental backstepping method decomposes the nonlinear system into several subsystems, and designs the control law for each subsystem based on the Lyapunov stability principle. It is suitable for uncertain nonlinear systems that can be linearized in a state, and has been repeatedly applied to the control research of small tilt-rotor aircraft [9,10,11]. The explicit model tracking control method is similar to the dynamic inversion in principle. It is generally composed of an explicit model, a controlled object, an inverse model, and a feedback compensation module. It is used in the design of helicopter stability augmentation systems in combination with flight quality requirements, and the quality design results are verified to be satisfactory by numerical simulation methods

[12] . There have been attempts to combine explicit model tracking control with nonlinear dynamic inversion control to design a control method suitable for multi-mode flight

[15] , and complete maneuvering mission experiments for quality evaluation, but no research has been conducted on the dynamic process of tilt transition. Summary of the invention

[0006] In order to solve the above technical problems, the present invention provides a design method for a longitudinal stabilization system of a tiltrotor aircraft transition section based on flight quality. In order to achieve smooth control of the tilt transition process, the present invention designs a stabilization control system with nonlinear time-varying characteristics, thereby reducing the pilot's workload and improving flight quality.

[0007] The main ideas of the technical solution design of the present invention are as follows:

[0008] At present, tiltrotor aircraft are typical variant aircraft. Their aerodynamic characteristics, center of gravity position, control mode, etc. change greatly during the full-mode flight process, and the flight quality requirements under the three working modes are also different. In order to achieve smooth control of the tilt-rotor transition process, the present invention designs a stabilization control system with nonlinear time-varying characteristics, so as to reduce the workload of the pilot and improve the flight quality. Aiming at the dynamic quality requirements and nonlinear characteristics of control during the transition process of the tilt-rotor aircraft, the present invention designs an inner loop control law based on the incremental backstepping method to adapt to the changes in aircraft characteristics; in order to meet the flight quality requirements under all modes, an ideal model under different working modes is designed; based on this, a longitudinal stabilization system suitable for the control of the transition section of the tilt-rotor aircraft is constructed. Taking the conversion from helicopter mode hovering to fixed-wing aircraft mode high-speed cruise as an example, the tilt-rotor transition path is designed, the control process of the transition section is simulated and analyzed through human-machine closed-loop simulation calculation and pilot-in-the-loop simulation experiment, and the effectiveness of the stabilization system design is verified based on time-varying flight quality evaluation.

[0009] In order to achieve the above object, a further technical solution of the present invention is:

[0010] The design method of the longitudinal stabilization system of the transition section of a tiltrotor aircraft based on flight quality is characterized by comprising the following steps:

[0011] S1: Analyze the longitudinal modal characteristics of the tiltrotor aircraft and summarize the requirements for stability augmentation control;

[0012] S2: Design an ideal model based on the flight quality requirements, design the inner loop control law based on the incremental backstepping method, and construct a longitudinal stability augmentation system based on this;

[0013] S3: Evaluate the stabilization effect of the longitudinal stabilization system on the dynamic process of tilt transition.

[0014] Preferably, the step S1 specifically includes:

[0015] S101: Tiltrotor Dynamics Model

[0016] According to the longitudinal motion behavior of the tilt-rotor aircraft, the longitudinal linear model of the tilt-rotor aircraft is extracted:

[0017]

[0018] Define Θ to represent the tilt angle of the rotor nacelle, Θ = 90 degrees for helicopter mode, Θ = 0 degrees for fixed-wing aircraft mode; V represents the forward flight speed; X = [Δθ; Δu; Δw; Δq], where Δθ is the pitch angle, Δu is the forward speed, Δw is the normal speed, and Δq is the pitch angular speed; δ lon Represents the longitudinal displacement of the joystick.

[0019] S102: Longitudinal modal characteristics analysis

[0020] S1021. Calculate the longitudinal modal characteristics under each state based on the nacelle tilt angle and forward flight speed;

[0021] S1022. Analyze the strong time-varying stabilization control requirements based on the short-period zero-pole distribution law.

[0022] Preferably, the driver model in step S2 is:

[0023]

[0024] Among them, K e represents the driver control gain, T L and T l Represent the driver lead and lag compensation time constants, τ e Represents delayed driver reaction time.

[0025] Preferably, the step S2 specifically includes:

[0026] S201: Designing an ideal model in three modes

[0027] S202: Design an inner loop control law based on the ideal model.

[0028] Preferably, the ideal models in the three modes are:

[0029] (3) Ideal model in fixed-wing aircraft mode (transfer function form of pitch angular velocity to longitudinal displacement of the joystick)

[0030]

[0031] (4) Ideal model in helicopter mode

[0032]

[0033] (3) Ideal model in the tilt transition mode. The tilt transition mode is divided into 30 degrees and 60 degrees tilt angles, and two ideal models of the tilt transition mode are obtained. Formula (3) is for 30 degrees, and formula (4) is for 60 degrees:

[0034]

[0035]

[0036] Preferably, the step S2 further comprises a step of evaluating the quality of the ideal model in the three modes.

[0037] The above ideal model should meet the first-level quality requirements of the CAP criterion and the attitude bandwidth criterion.

[0038] Preferably, the stability augmentation system in step S202 replaces the outer loop control law in the traditional incremental backstepping method with the driver and the ideal model, and the outer loop state error is obtained after the driver and the ideal model to obtain the inner loop desired command q r ; Based on the ideal model, the inner loop expected instruction q can be obtained r ;

[0039] Then the inner loop state error e is:

[0040] e k =q k -q r k Formula (6)

[0041] In formula (6), q is the actual pitch velocity of the aircraft; q r is the pitch angular velocity command expected by the inner loop; the upper right subscript k of each variable represents the kth step, corresponding to the current moment, and its meaning in the following formulas is the same.

[0042] The inner loop control Lyapunov function is established as:

[0043]

[0044] Taking its derivative we get:

[0045]

[0046] In formula (8), is the derivative of the Lyapunov function V, is the derivative of the inner loop state error, that is, the rate of change of the pitch angular velocity tracking error; is the pitch acceleration of the aircraft; is the desired pitch acceleration of the inner loop.

[0047] Substituting the above formula and combining it with Lyapunov's stability condition, we can derive:

[0048]

[0049] In formula (9), ΔU is the increment of the control amount applied to the aircraft by the stability augmentation system; q e is the tracking error of the aircraft's pitch angular velocity; c is the parameter of the stability augmentation control system to be designed, which can be taken as a constant; B is the control matrix in the aircraft linear model; the upper right subscript k-1 indicates the k-1th step, corresponding to the previous moment, and its meaning in the following formulas is the same.

[0050] Then the inner loop control law is:

[0051]

[0052] In formula (10), U is the control variable applied to the aircraft by the stability augmentation system.

[0053] Preferably, the step S3 specifically includes:

[0054] S301, selecting a driver model for numerical simulation;

[0055] S302: Establishing performance indicators J that take into account the time to complete the tilt, attitude change, and driver control load route Get the optimal tilt path.

[0056] As a preference, the performance index J route Expressed as follows:

[0057]

[0058] Where: t0——Tilt start time; t f ——Tilt end time; q——Pitch angular velocity, unit: degree / second; θ——Pitch attitude angle, unit: degree; ——nacelle tilting angular velocity, unit: degree / second; ——Longitudinal change rate of joystick, unit: cm / s; X max ——The maximum value of X during the tilting process.

[0059] Compared with the prior art, the present invention has the following beneficial effects:

[0060] The present invention aims at the time-varying characteristics of the tiltrotor during the mode conversion flight process, and combines the different requirements of the three working modes for the flight quality, and designs a longitudinal stabilization system suitable for the transition section control of the tiltrotor to meet the flight quality requirements in all modes. The present invention also conducts simulation and related experiments on the flight mission of the tiltrotor from hovering in helicopter mode to cruising in fixed-wing aircraft mode, and has at least the following contributions:

[0061] (1) A quality-based time-varying controller design method is proposed. The present invention combines the flight quality design requirements with the nonlinear controller, and designs a longitudinal stabilization system in the form of an inner loop compensation stabilization and an outer loop adaptation to the quality requirements. It is then applied to the transition section control of a tiltrotor aircraft with dynamic quality requirements and time-varying aircraft characteristics to improve its flight quality issues.

[0062] (2) Based on the flight quality requirements, an ideal model of the tilt transition section is given, and the ideal model is evaluated using the attitude bandwidth criterion in the helicopter quality specification and the CAP criterion in the fixed-wing aircraft quality specification. The results all meet the first-level quality specifications.

[0063] Based on the idea of ​​incremental backstepping, the outer loop control law is replaced by the driver and the ideal model, and the inner loop control law is redesigned. As a result, the unknown parameters of the control law are reduced, and the dependence on the outer loop tracking error is eliminated, which can be used for longitudinal stability augmentation design. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] Figure 1a It is a possible flight state of the tilt transition process of the present invention;

[0065] Figure 1b A diagram showing the change in the short-period zero-pole distribution during the tilt transition process of the present invention;

[0066] Figure 2 This is a design diagram of the quality-based longitudinal stabilization system of the present invention;

[0067] Figure 3a This is a diagram showing the design result of the ideal model tested according to the attitude bandwidth criterion of the present invention;

[0068] Figure 3b The ideal model design result diagram is tested according to the CAP criteria of the present invention;

[0069] Figure 4 A design diagram comparing the control structure of the conventional incremental backstepping control of the present invention and the stability augmentation system of the present invention;

[0070] Figure 5 This is the structure diagram of the traditional incremental backstepping control method;

[0071] Figure 6 The ideal model and inner loop control law structure diagram of the longitudinal stabilization system of the present invention;

[0072] Figure 7a The optimal path diagram of the tilt transition of the numerical simulation tilt transition flight process of the present invention;

[0073] Figure 7b It is a timing instruction diagram of the rotor nacelle tilt angle for numerically simulating the tilt transition flight process of the present invention;

[0074] Figure 8 A pitch angle curve diagram of a dynamic simulation of a numerical simulation of a tilt transition flight process according to the present invention;

[0075] Fig. 9 This is a structural diagram of the driver-in-the-loop simulation experiment of the present invention;

[0076] Fig.10a It is a pitch angle tracking curve diagram of the in-loop simulation experiment of the present invention;

[0077] Fig.10b It is a weighted phase lag curve diagram of the in-loop simulation experiment of the present invention;

[0078] Fig.11Comparison of the driver's lever force-time spectrum before and after warming up;

[0079] Fig.12 A comparison chart of Klyde's time-varying flight quality criteria during the tilt transition process. DETAILED DESCRIPTION

[0080] In order to enable those skilled in the art to better understand the solutions of the present invention, the technical solutions in the embodiments of the present invention are described clearly and completely below. Obviously, the described embodiments are only embodiments of a part of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work should fall within the scope of protection of the present invention.

[0081] Example 1

[0082] (1) Establishment and analysis of tiltrotor aircraft related models

[0083] 1.1 Tilt-rotor aircraft dynamics model

[0084] The tilt-rotor aircraft contains the control mechanisms of a helicopter and a fixed-wing aircraft. Its dynamic model includes the flight dynamics model of the tilt-rotor aircraft in all modes and the hybrid control model. The direct control variables of the pilot are the collective pitch / throttle lever, the longitudinal displacement of the joystick, the lateral displacement of the joystick, and the pedal displacement. After the hybrid control strategy allocation reference

[13] , the collective pitch, longitudinal and lateral cyclic pitch change, and three aerodynamic surfaces of the left and right rotors are controlled. The hybrid control strategy allocation coefficient changes with the tilt angle of the rotor nacelle and the flight speed. For the problem of the tilt transition process, the tilt angle of the rotor nacelle is assumed to change from 90 degrees to 0 degrees, and the forward flight speed increases from 0m / s to 140m / s. The linearized model of the tilt-rotor aircraft can be obtained by linearizing the tilt-rotor aircraft model in different states.

[0085] The flight mode conversion of the tiltrotor aircraft is achieved by tilting the rotor nacelles on both sides. In this process, the lift component of the rotor thrust will decrease as the nacelles tilt forward, and it is necessary to increase the angle of attack to increase the aerodynamic force to compensate for the lack of lift; and the transverse double rotor layout of the tiltrotor aircraft provides greater lateral heading stability. During the tilt transition process, the control behavior of the pilot is mainly based on the movement of the longitudinal channel. Therefore, the present invention conducts research on the longitudinal movement of the tiltrotor aircraft and extracts the longitudinal linearization model of the tiltrotor aircraft:

[0086]

[0087] Definition: =90deg is helicopter mode, =0deg is fixed-wing aircraft mode.

[0088] 1.2 Analysis of longitudinal modal characteristics

[0089] Based on the nacelle tilt angle and forward flight speed, Figure 1a , 1b The forward flight speed and tilt angle of the flight state are given, which basically covers the possible flight states of the tilt transition stage flight process, and the longitudinal modal characteristics under each state are calculated. The short-period zero-pole distribution law is as follows: Figure 2 As shown in the figure, static instability occurs at low speed in helicopter mode and gradually tends to be stable with the increase of speed. The inclination angle of the nacelle has a greater influence on the real part, and the influence of speed on the modal characteristics gradually decreases as the flight mode is converted to the fixed-wing aircraft mode.

[0090] (2) Design process of longitudinal stability augmentation system based on flight quality

[0091] The longitudinal stabilization system based on flight quality consists of two parts: an ideal model and an inner-loop control law. The inner-loop control law is required to adapt to the strong time-varying characteristics of the aircraft characteristics. Guided by the design concept of the incremental backstepping method, the inner-loop control law follows the changes in the aircraft characteristics and always compensates the aircraft characteristics as an approximate pure integral link, thereby approximating the time-varying characteristics of the controlled object to a time-invariant controlled object. The ideal model of the outer loop is then used to achieve the flight quality requirements under different modes. The structure of the longitudinal stabilization system including the ideal model and the inner-loop control law is as follows: Figure 2 shown.

[0092] 2.1 Ideal Model Based on Quality

[0093] like Figure 2 As shown, the task of the ideal model is to generate a pitch rate instruction that meets the quality specification. The present invention solves the ideal model in each mode by deriving the relationship between the first-level quality requirement and the ideal model.

[0094] In the fixed-wing aircraft mode, the tilt-rotor aircraft controls the pitch axis motion of the aircraft by manipulating the elevator deflection angle. For the tilt-rotor aircraft after stabilization, the control expectation parameter CAP and short-period damping ratio defined in MIL-STD-1797A can be used to evaluate the short-period motion characteristics. The ideal model form of the fixed-wing aircraft mode is the short-period modal motion transfer function form. According to the first-level quality requirements in the criteria, after determining the parameters, we get:

[0095]

[0096] In helicopter mode, the tilt-rotor aircraft controls pitch by manipulating the longitudinal cyclic pitch of the two rotors. By selecting the appropriate response type, the system bandwidth and phase delay are calculated according to the frequency characteristics of its pitch attitude response to obtain the quality level. The ideal model form of the helicopter mode is RCAH, which generally adopts a first-order system. According to the first-level quality requirements for attitude bandwidth-delay requirements in ADS-33E-PRF, after determining the ideal model parameters, we get:

[0097]

[0098] From the analysis of the longitudinal modal characteristics of the tilt-rotor aircraft in all modes, it can be seen that the dynamic characteristics of the tilt transition mode are more similar to those of fixed-wing aircraft. Therefore, the ideal model form of the tilt transition mode is consistent with the fixed-wing aircraft mode. At the same time, the mode transition smoothness requirement should be met. The tilt transition mode is divided into 30deg and 60deg tilt angles, and two ideal models of the tilt transition mode are obtained. Equation (3) is 30 degrees, and Equation (4) is 60 degrees:

[0099]

[0100]

[0101] like Figure 3a , 3b As shown, the ideal model design results are tested according to the CAP criterion and the attitude bandwidth criterion, and both meet the first-level quality requirements.

[0102] 2.2 Inner loop control law based on ideal model

[0103] The design of the inner loop control law is an important prerequisite for the outer loop ideal model to work. The inner loop control law needs to complete the real-time tracking compensation of the controlled object to ensure the realization of the pure integral link. The inner loop control law has the characteristics of nonlinearity. The present invention refers to the traditional incremental backstepping method and uses Lyapunov's second method to derive the inner loop control law. Compared with the incremental backstepping method, it solves the feedback of the outer loop attitude signal and is suitable for the stabilization of manned aircraft.

[0104] The stability augmentation system designed by the present invention replaces the outer loop control law in the traditional incremental backstepping method with the driver model and the ideal model. The outer loop state error is obtained after the driver model and the ideal model to obtain the inner loop desired command q r ;

[0105] Then the inner loop state error e is:

[0106] e k =q k -q r k Formula (6)

[0107] In formula (6), q is the actual pitch velocity of the aircraft; q r is the pitch angular velocity command expected by the inner loop; the upper right subscript k of each variable represents the kth step, corresponding to the current moment, and its meaning in the following formulas is the same.

[0108] The inner loop control Lyapunov function is established as:

[0109]

[0110] Taking its derivative, we get:

[0111]

[0112] In formula (8), is the derivative of the Lyapunov function V, is the derivative of the inner loop state error, that is, the rate of change of the pitch angular velocity tracking error; is the pitch acceleration of the aircraft; is the desired pitch acceleration of the inner loop.

[0113] Substituting the above formula and combining it with Lyapunov's stability condition, we can derive:

[0114]

[0115] In formula (9), ΔU is the increment of the control amount applied to the aircraft by the stability augmentation system; q e is the tracking error of the aircraft’s pitch angular velocity; c is the parameter of the stability augmentation control system to be designed, which can be a normal number; B is the control matrix in the aircraft linearization model; the upper right subscript k-1 indicates the k-1th step, corresponding to the previous moment, and its meaning in the following formulas is the same.

[0116] Then the inner loop control law is:

[0117]

[0118] In formula (10), U is the control variable applied to the aircraft by the stability augmentation system.

[0119] Figure 5 ,6 Compared with the traditional incremental backstepping method, the control law parameters designed by the present invention have two c, B k-1 , reducing an unknown control parameter. As shown in the figure, the stabilization system of the present invention relies on the driver to control the tracking outer loop command θ r , get the instruction q of the inner loop control law generated by the driver's control signal manager model r The inner loop control law of the incremental backstepping method will introduce the error term of the outer loop Lyapunov function, so the outer loop tracking error θ e The stabilization system of the present invention can achieve the stabilization effect on manned aircraft and has a wider application range.

[0120] Example 2: Numerical simulation of tilt transition flight process

[0121] (1) In order to evaluate the stabilization effect of the stability augmentation system on the dynamic process of tilt transition, a tilt transition path was designed to perform numerical simulation of the tilt transition flight process, taking the transition of the tilt rotor aircraft from helicopter mode hovering to fixed-wing aircraft mode high-speed cruising flight as the background. Numerical simulation requires the pilot model to build a human-machine closed-loop system to verify the design effect of the longitudinal stability augmentation control system. The pilot model selects the McRuer pilot model, which is in the form of formula (11). Considering the crossover frequency characteristics and phase margin requirements of the human-machine open-loop system, the reference

[19] selects the McRuer pilot model parameters under different working modes, and the results are shown in Table 1.

[0122]

[0123] Ke represents the driver control gain, TL and TI represent the driver lead and lag compensation time constants, τ e Represents delayed driver reaction time.

[0124] Table 1 Driver model parameters

[0125]

[0126] (2) Establish performance index J that takes into account the time to complete the tilt, attitude change, and pilot control load route Reference

[13] obtains the optimal tilt path as Figure 7a As shown, in reference 20, the speed of the tilting process is set to 2.1m / s2 uniformly accelerated forward flight, and the timing instructions for the tilting angle are obtained as follows Figure 7b According to the timing instructions of the nacelle inclination angle and the ideal model, the mode switching is set at 9.6 seconds, 27.5 seconds, and 46 seconds to achieve a gradual transition from helicopter mode to fixed-wing aircraft mode.

[0127]

[0128] Where: t0——Tilt start time; t f ——Tilt end time; q——Pitch angular velocity, unit: degree / second; θ——Pitch attitude angle, unit: degree; ——nacelle tilting angular velocity, unit: degree / second; ——

[0129] Joystick longitudinal change rate, unit: cm / s; X max ——The maximum value of X during the tilting process.

[0130] Tracking curve from pitch angle command Figure 8It can be seen that the pilot model has smooth control during the tilt transition process. The time-varying characteristics of the tiltrotor aircraft's body configuration and the ideal model switching have no effect on the tracking effect of the pitch angle signal. The longitudinal stabilization system has a delay of about 0.6s more than the incremental backstepping method, and part of this delay comes from the pilot model.

[0131] Example 3: Quality evaluation based on driver-in-the-loop simulation experiment

[0132] Design bandwidth instructions to stimulate the driver's control behavior and analyze human-machine quality. Fig. 9 As shown in the figure, a pilot-in-the-loop simulation experiment is carried out based on an engineering fixed-base flight simulator. The pilot receives the real-time tracking error signal of the pitch angle through the visual system. The push-pull joystick generates a signal that is solved by the simulation computer to solve the dynamic model and output the pilot's control signal during the tilt-rotor transition flight. The control behavior of the tilt-rotor aircraft before and after the pilot controls the stabilization is compared, and the quality is evaluated using the time-frequency method and Klyde's time-varying flight quality criterion reference

[17] .

[0133] like Fig.10a , the pitch angle tracking accuracy of the pilot when controlling the stabilized aircraft is significantly improved, and the average tracking error is reduced from 1.22 degrees before stabilization to 0.72 degrees, a reduction of 41%. Before stabilization, the pilot felt strenuous in helicopter mode, the controlled object responded slowly, and could not track the pitch angle command; in fixed-wing aircraft mode, the controlled object responded too sensitively, the pilot needed to operate frequently, and the tracking effect was poor. After stabilization, the pilot can feel the changes in the characteristics of the controlled object during operation, but the control feeling and pitch angle command tracking effect are good throughout the whole process.

[0134] Based on Klyde's criterion, the curve is drawn as follows Fig.12 As shown, the ordinate of the standard diagram is the peak value P of the spectrum when the driver controls the lever force max (t), reflects the driver's control load, and the horizontal axis is the average phase lag weighted by the energy spectrum It reflects the dynamic characteristics of the tiltrotor aircraft. The time domain curve of the average phase lag is as follows: Fig.10b , which corresponds to the three flight modes of the tiltrotor aircraft. The calculation method of this parameter is derived from the reference of the tiltrotor aircraft itself

[17] , so there is little change before and after stabilization. Fig.11 The time-frequency spectrum of the pilot's joystick force shows that the stability augmentation system mainly improves the flight quality by reducing the pilot's joystick force. By comparing the Klyde time-varying flight quality criterion results before and after the stability augmentation (such as Fig.12 The stability augmentation system effectively changes the Type III PIO tendency in helicopter mode from being present to being absent.

[0135] result:

[0136] The dynamic numerical simulation results of tilt transition show that when the pilot is operating a tiltrotor aircraft with a longitudinal stabilization system to complete a tilt transition flight, the ideal model is switched three times during the flight to adapt to the flight quality requirements under different working modes, without affecting the tracking effect of the pitch angle command. The inner loop control law adapts to the time-varying characteristics of the aircraft characteristics and the nonlinear changes in the pitch angle velocity command caused by the switching of the ideal model, so that the pitch angle command is always tracked well throughout the whole process.

[0137] The pilot-in-the-loop test results before and after longitudinal stabilization show that the longitudinal stabilization system can effectively reduce the pilot control load in all modes, especially in helicopter mode. The helicopter mode has a Type III PIO trend before stabilization, and the peak value of the joystick force is significantly reduced after stabilization, and the helicopter mode no longer has a Type III PIO trend.

[0138] The following are some prior art references used in the design of this invention:

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[0160] Utilizing the technical solution of the present invention, or those skilled in the art designing similar technical solutions inspired by the technical solution of the present invention to achieve the above-mentioned technical effects, all fall within the protection scope of the present invention.

Claims

1. A design method for a longitudinal stabilization system of a tiltrotor aircraft transition section based on flight quality, characterized in that: The following steps are involved: S1: Analyze the longitudinal modal characteristics of the tiltrotor aircraft and summarize the requirements for stability augmentation control; S2: Design an ideal model based on the flight quality requirements, design the inner loop control law based on the incremental backstepping method, and construct a longitudinal stability augmentation system based on this; S3: Evaluate the stabilization effect of the longitudinal stabilization system on the dynamic process of tilt transition; The driver model in step S2 is: Among them, K e represents the driver control gain, T L and T l Represent the driver lead and lag compensation time constants, τ e represents the delayed driver reaction time; The step S2 specifically includes: S201: Designing an ideal model in three modes S202: Design inner loop control law based on ideal model; The ideal models under the three modes are: (1) Ideal model in fixed-wing aircraft mode (2) Ideal model in helicopter mode (3) Ideal model in the tilt transition mode. The tilt transition mode is divided into 30 degrees and 60 degrees tilt angles, and two ideal models of the tilt transition mode are obtained. Formula (3) is 30 degrees, and formula (4) is 60 degrees: The step S2 also includes the step of evaluating the quality of the ideal model under the three modes; The stability augmentation system in step S202 replaces the outer loop control law in the traditional incremental backstepping method with the driver and the ideal model. The outer loop state error is obtained after the driver and the ideal model to obtain the inner loop desired command q r ; Then the inner loop state error e is: e k =q k -q r k Formula (6) In formula (6), q is the actual pitch velocity of the aircraft; q r is the pitch velocity command expected by the inner loop; the upper right subscript k of each variable indicates the kth step, corresponding to the current moment, and its meaning in the following formulas is the same; The inner loop control Lyapunov function is established as: Taking its derivative we get: In formula (8), is the derivative of the Lyapunov function V, is the derivative of the inner loop state error, that is, the rate of change of the pitch angular velocity tracking error; is the pitch acceleration of the aircraft; is the desired pitch acceleration of the inner loop; Substituting the above formula and combining it with Lyapunov's stability condition, we can derive: In formula (9), ΔU is the increment of the control amount applied to the aircraft by the stability augmentation system; q e is the tracking error of the aircraft pitch angular velocity; c is the parameter of the stability augmentation control system to be designed, which can be a normal number; B is the control matrix in the aircraft linearization model; the upper right subscript k-1 indicates the k-1th step, corresponding to the previous moment, and its meaning in the following formulas is the same; Then the inner loop control law is: In formula (10), U is the control variable applied to the aircraft by the stability augmentation system.

2. The design method of the tiltrotor aircraft transition section longitudinal stabilization system based on flight quality according to claim 1 is characterized in that: The step S1 specifically includes: S101: Tiltrotor Dynamics Model According to the longitudinal motion characteristics of the tilt-rotor aircraft, the longitudinal linear model of the tilt-rotor aircraft is extracted: Define Θ to represent the tilt angle of the rotor nacelle, Θ = 90 degrees for helicopter mode, Θ = 0 degrees for fixed-wing aircraft mode; V represents the forward flight speed; X = [Δθ; Δu; Δw; Δq], where Δθ is the pitch angle, Δu is the forward speed, Δw is the normal speed, and Δq is the pitch angular speed; δ lon represents the longitudinal displacement of the joystick; S102: Longitudinal modal characteristics analysis S1021. Calculate the longitudinal modal characteristics under each state based on the nacelle tilt angle and forward flight speed; S1022. Analyze the strong time-varying stabilization control requirements based on the short-period zero-pole distribution law.

3. The design method of the tiltrotor aircraft transition section longitudinal stabilization system based on flight quality according to claim 1, characterized in that: The step S3 specifically includes: S301, selecting a driver model for numerical simulation; S302: Establishing performance indicators J that take into account the time to complete the tilt, attitude change, and driver control load route Get the optimal tilt path.

4. The design method of the tiltrotor aircraft transition section longitudinal stabilization system based on flight quality according to claim 3 is characterized in that: The performance index J route Expressed as follows: Where: t0——Tilt start time; t f ——Tilt end time; q——Pitch angular velocity, unit: degree / second; θ——Pitch attitude angle, unit: degree; ——nacelle tilting angular velocity, unit: degree / second; ——Longitudinal change rate of joystick, unit: cm / s; X max ——The maximum value of X during the tilting process.

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