An aircraft servo-elastic dynamics model correction method considering actuator gap

Abstract

The application provides a kind of aircraft servo elastic dynamics model correction method considering operating surface gap, first aircraft structure is divided into substructure, and the aircraft dynamics model containing aileron gap nonlinearity is established according to substructure branch modal;And additional inertia force caused by aircraft operating aileron deflection is introduced in dynamics model, and the aircraft open-loop servo elastic motion equation containing aileron gap nonlinearity under independent generalized coordinate is obtained;According to sensor output equation, rudder transfer function and flight control system equation, the aircraft nonlinear servo elastic closed-loop system model containing aileron gap nonlinearity is established, and the model is corrected based on ground vibration test, rudder frequency response test and gap measurement test result, and the stability of the model is judged based on simulation analysis result and open-loop frequency response test result and is further corrected, which improves the accuracy of gap nonlinearity aircraft aerodynamic servo elastic characteristic evaluation, and is helpful to guide aircraft servo elastic ground test engineering practice.

Description

Technical Field

[0001] This invention belongs to the field of aeroelasticity and ground testing technology of aircraft, specifically relating to a method for correcting an aircraft servo-elastic dynamics model that considers control surface clearance. Background Technology

[0002] In recent years, with the vigorous development of my country's aerospace industry, the aeroelasticity of aircraft has received great attention, greatly promoting the research and development of aeroelastic mechanics for aircraft in my country. Among these, the aerodynamic servo characteristic problem is a discipline that considers the coupling between the aircraft control system and structural vibration and unsteady aerodynamic forces. During flight, the elastic vibration signals of the aircraft structure and the rigid body motion signals are received by sensors together. After processing by the flight control system, these signals drive the control surfaces to deflect. The control surfaces, on top of the low-frequency deflection, undergo additional high-frequency deflection. This high-frequency unsteady control surface excitation causes further elastic vibration of the structure, resulting in aerodynamic / structural / control coupling problems in the aircraft system. To ensure that the aircraft does not exhibit aerodynamic servoelastic instability within its flight range, modal tests and servoelastic ground tests are required before the maiden flight to test the dynamic coupling characteristics between the control system and the structure. These tests include open-loop frequency response tests and closed-loop stability tests, aiming to provide accurate mathematical models for the aerodynamic servoelastic analysis of the aircraft.

[0003] However, while modern aerospace vehicles possess higher flight speeds and stronger maneuverability, the aero-servoelastic problems they involve are becoming increasingly complex, and the resulting nonlinearities are becoming more pronounced. Among these, the control surfaces are a relatively weak link, requiring deflection to provide the necessary trim forces or moments for the aircraft; therefore, the control surface stiffness is relatively small compared to the structural stiffness. Due to factors such as manufacturing deviations, assembly errors, and wear during operation, gap nonlinearity is one of the most common concentrated structural nonlinearities in aircraft. Gap nonlinearities exist at the control surface pivot points and in the control system stiffness, potentially altering the system's dynamic characteristics and affecting the aero-servoelastic properties of the aircraft. Traditional servoelasticity tests do not consider the nonlinearity of the control surfaces, leading to inaccurate aero-servoelastic stability assessments of the aircraft. Summary of the Invention

[0004] The purpose of this invention is to solve the problem that the existing servoelasticity test does not consider the control surface clearance when constructing the dynamic model of the aircraft aero-servoelastic system, resulting in poor accuracy of the aircraft aero-servoelastic stability assessment results. The invention provides a method for correcting the aircraft servoelastic dynamic model that considers the control surface clearance.

[0005] To achieve the above objectives, the technical solution provided by this invention is:

[0006] A method for correcting an aircraft servo-elastic dynamics model that considers control surface clearance includes the following steps:

[0007] Step 1: Divide the aircraft structure containing aileron clearance nonlinearity into several substructures, and use several nodes of the nonlinear connection region on the aircraft structure as interfaces to discretize each substructure to obtain the corresponding branch modes of each substructure; based on the branch modes of each substructure, establish an aircraft dynamics model containing aileron clearance nonlinearity.

[0008] Step 2: In the nonlinear aircraft dynamics model with aileron clearance established in Step 1, the additional inertial force caused by the aileron deflection is introduced to obtain the motion equations of the open-loop servo-elastic system of the aircraft with aileron clearance nonlinearity in independent generalized coordinates:

[0009]

[0010] In the formula: This is the quality matrix; Here is the stiffness matrix; For independent generalized coordinates; It is a non-linear vector; The additional inertial force at the structural nodes caused by the longitudinal rudder surface deflection motion degree of freedom, and , For aileron deflection; The cross mass array, and the aileron deflection mode shape Regarding this, the aileron deflection mode shape is approximately expressed as: ,in For the rudder axis heading coordinates, Here are the heading coordinates of each node; define the direction along the airflow as... The axis is positive, and the trailing edge of the aileron is deflected downwards to be positive;

[0011] Step 3: Based on the sensor output equation, the state space equation of the open-loop nonlinear servo-elastic motion equation, and the state space equation of the flight control system, establish the nonlinear servo-elastic closed-loop system equation of the aircraft containing aileron clearance nonlinearity, and construct the nonlinear servo-elastic closed-loop system model of the aircraft containing aileron clearance nonlinearity.

[0012] Step 4: Correct the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity established in Step 3 to obtain the corrected nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity; wherein, the correction parameters include the aircraft structural dynamic characteristics, servo transmission characteristics, clearance nonlinearity parameters and flight control system parameters.

[0013] Step 5: Perform servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity after modification, and conduct servo-elasticity test and simulation analysis. Obtain the test results and simulation analysis results of the system response spectrum. Compare the simulation analysis results with the test results. If the simulation analysis results are consistent with the test results, it indicates that the modified nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity meets the stability requirements. Otherwise, return to step 4 to modify the model again.

[0014] Furthermore, in step 3, the process of establishing the nonlinear servo-elastic closed-loop system equations for the aircraft containing aileron clearance nonlinearity includes:

[0015] Step 3.1: Establish the response model of the sensor installation location in the physical coordinate system and obtain the sensor output equation. In the formula, The state vector includes modal coordinates, control surface deflection states, and flight control system state variables. The modal matrix for the sensor mounting location;

[0016] Step 3.2: First, establish the corresponding servo motor dynamic state-space equations based on the servo motor transfer function.

[0017] ,in, The state vector of the manipulation surface. The control surface deflection command is fed back to the servo motor from the flight control system. and These are the system matrix and control matrix of the servo motor, respectively;

[0018] Then, combining the sensor output equation from step 3.1 and the motion equations of the open-loop servo-elastic system of the aircraft in independent generalized coordinates with aileron clearance nonlinearity obtained in step 2, the state-space equation form of the open-loop nonlinear servo-elastic motion equation is obtained. , and These are the system matrix and control matrix of the servo-elastic system, respectively. It is a non-linear vector;

[0019] Step 3.3: Establish the flight control system model. The state-space equation of the flight control system is as follows: And set the gain feedback relationship between the servo input and the flight control system output as follows: ;

[0020] in, This is the state vector of the flight control system. The sensor output vector, , , and These are the system matrix, control matrix, output matrix, and direct transfer matrix of the flight control system. This is the gain matrix;

[0021] Step 3.4: Based on the results of Steps 3.1-3.3, obtain the nonlinear servo-elastic closed-loop system equations of the aircraft with aileron clearance nonlinearity, and construct the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity.

[0022] Furthermore, in step 4, the process of correcting the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity established in step 3 includes:

[0023] Step 4.1: Conduct ground vibration modal tests on the aircraft, obtain modal test results, including information on the modal frequencies and mode shapes of each order; and modify the material properties or structural geometric parameters of the aircraft dynamic model with aileron clearance nonlinearity constructed in Step 1 based on the modal test results, so as to change the stiffness and mass characteristics of the aircraft dynamic model and ensure that the modal frequencies and mode shapes are consistent with the test model.

[0024] Step 4.2: Perform servo-rudder deflection frequency response test to obtain aileron deflection signal and corresponding spectrum curve. Fit the amplitude frequency and phase frequency transfer characteristic curve function of the nonlinear servo elastic closed-loop system of the aircraft containing aileron clearance nonlinearity. Correct the servo transfer function in step 3.2 using the transfer characteristic curve function.

[0025] Step 4.3: The aileron is loaded and unloaded in both directions using the "load-displacement curve" gap measurement method to obtain the aileron displacement under the corresponding load and the aileron loading-displacement curve; the gap size under zero load is obtained based on the loading-displacement curve; and the control surface deflection stiffness is corrected based on the gap size to accurately characterize the nonlinear stiffness of the control surface gap.

[0026] Furthermore, in step 4.2, the specific process of performing the servo-rudder deflection frequency response test to obtain the aileron deflection signal and the corresponding spectrum curve is as follows:

[0027] Step 4.2.1: Build a test system, including the Quanser hardware-in-the-loop simulation system, which has a signal PWM generator, a modulation and demodulation module, and an ADC sampling module;

[0028] A laser sensor is installed in the nonlinear servo-elastic closed-loop system model of the aircraft containing aileron clearance nonlinearity. The laser sensor is connected to the ADC sampling module to collect aileron displacement response data.

[0029] Step 4.2.2: The Quanser hardware-in-the-loop simulation system modulates and demodulates the input sweep frequency signal and then inputs it to the PWM generator. The demodulated signal satisfies the PWM wave duty cycle required by the PWM generator.

[0030] The PWM generator outputs a PWM square wave control signal to the servo motor according to the PWM wave duty cycle; wherein the high-level duration of the PWM square wave control signal corresponds to the servo motor deflection angle, and the high-level duration is linearly related to the servo motor deflection angle.

[0031] Step 4.2.3: The servo executes the deflection command according to the received PWM square wave control signal, driving the aileron to deflect and exciting the aileron to vibrate;

[0032] Step 4.2.4 The laser sensor collects the aileron displacement response data and transmits it to the AD sampling module of the Quanser hardware-in-the-loop simulation system;

[0033] Step 4.2.5: Obtain the aileron frequency response function based on the collected displacement response data and the frequency sweep signal.

[0034] Furthermore, in step 5, the servo elasticity test includes an open-loop frequency response test. Based on the open-loop frequency response test, the Bode plot and Nyquist plot of the system are obtained to acquire the open-loop amplitude-phase frequency characteristic curve and the open-loop stability margin of the system.

[0035] Furthermore, step 5 also includes a step of performing a closed-loop servo elasticity test; the closed-loop servo elasticity test is used to obtain the closed-loop stability margin of the system.

[0036] Furthermore, step 1 is implemented based on the Craig-Bampton fixed interface modal synthesis method.

[0037] The advantages of this invention are:

[0038] 1. In the model correction method of this invention, the branch modes of the substructure are obtained based on the Craig-Bampton fixed interface modal synthesis method, and an aircraft dynamic model containing aileron clearance nonlinearity is established based on the branch modes; and through sensor output equations, servo transfer functions and flight control system equations, a nonlinear servo-elastic closed-loop system model of the aircraft containing aileron clearance nonlinearity is constructed, realizing accurate modeling of the aircraft dynamic model containing control surface clearance nonlinearity, and making the nonlinear stiffness of the control surfaces explicitly expressed in the system motion equations, which can accurately characterize the nonlinear stiffness characteristics of the system.

[0039] 2. In the model correction method of this invention, the nonlinear servo-elastic closed-loop system model of the aircraft containing aileron clearance nonlinearity is corrected by the results of aircraft ground vibration modal test, servo frequency response test, and clearance quantity test. This makes the constructed model more reliable and improves the accuracy of evaluating the aerodynamic servo-elastic characteristics of the clearance nonlinear aircraft using the constructed model. It also helps to guide the engineering practice of aircraft servo-elastic ground test.

[0040] Additional aspects and advantages of the invention will be set forth in part in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. Attached Figure Description

[0041] The above and / or additional aspects and advantages of the present invention will become apparent and readily understood from the description of the embodiments taken in conjunction with the following drawings, in which:

[0042] Figure 1 This is a flowchart of the method for correcting the dynamic model of aircraft servo characteristics considering control surface clearance according to the present invention.

[0043] Figure 2 This is a schematic diagram of the aircraft structure with nonlinear aileron clearance in this invention;

[0044] Figure 3 This is a block diagram of the nonlinear servo-elastic closed-loop system of an aircraft with aileron clearance nonlinearity in this invention;

[0045] Figure 4 This is a schematic diagram showing the relationship between the duration of the PWM high level and the angle in the servo drive control method of the present invention;

[0046] Figure 5 This is a flowchart of the frequency response function experiment in this invention;

[0047] Figure 6 This is a spectrum analysis diagram of the 5° sweep frequency of the aircraft wing servo motor in this embodiment of the invention;

[0048] Figure 7 These are the amplitude and phase frequency results of the servo motor at a 5° sweep frequency obtained using polynomial fitting, where 7a is the amplitude frequency result and 7b is the phase frequency result.

[0049] Figure 8 This is a load-displacement curve of the aileron clearance under actual conditions;

[0050] Figure 9 This is a Bode plot of the system under the 5° sweep amplitude of the servo motor in this embodiment of the invention, where 9a is the amplitude plot and 9b is the phase plot;

[0051] Figure 10This is the Nyquist plot of the system under the servo motor's 5° sweep amplitude in this embodiment of the invention;

[0052] Figure 11 This is a schematic diagram of a closed-loop coupling test.

[0053] Figure 12 This is a diagram of the aileron deflection signal when the pitch channel feedback gain K=225 in this embodiment of the invention;

[0054] Figure 13 This is a diagram of the aileron deflection signal when the pitch channel feedback gain K=450 in an embodiment of the present invention;

[0055] Figure 14 This is a comparison chart of the experimental results and the system response spectrum after correction by the servo elastic dynamics model of this invention.

[0056] Explanation of reference numerals in the attached figures: -Main wing surfaces of the fuselage -Left Auxiliary Wing -Right aileron, 1-Hinge. Detailed Implementation

[0057] The embodiments of the present invention are described in detail below. These embodiments are exemplary and intended to explain the present invention, and should not be construed as limiting the present invention.

[0058] Reference Figure 1 This invention provides a method for correcting an aircraft servo-elastic dynamics model that considers control surface clearance. The specific implementation steps are as follows:

[0059] Step 1: For aircraft structures with aileron clearance nonlinearity, based on the Craig-Bampton fixed-interface modal synthesis method, the nodes related to the nonlinear connection region are used as interfaces to divide the aircraft into several substructures. By discretizing each substructure, the branch modes of the substructure are obtained; and a dynamic model of the aircraft with aileron clearance nonlinearity is established based on the branch modes.

[0060] Reference Figure 2 The aircraft structure containing aileron clearance nonlinearity is divided into three substructures: the main fuselage surface, the left aileron, and the right aileron structure. These are respectively represented by... , and It indicates. Among them, the left aide-de-camp and right aileron Each with the main wing surface of the fuselage They are connected by hinge 1.

[0061] Based on the Craig-Bampton (CB) fixed-interface modal synthesis method, the hinge nodes related to the nonlinear connection region are used as interfaces. Each substructure is discretized to obtain the branch modes of the substructure. Among them, the branch modes include the fixed-interface master mode. and constrained modes .

[0062] Using the fixed interface principal modes and constraint modes of each substructure obtained above, a set of hypothetical modes is formed. The model is then subjected to modal coordinate transformation and degree of freedom reduction on the assumed mode set, and then the substructures are grouped together to obtain the modal coordinates of the entire aircraft structure.

[0063] Based on the connection conditions of each substructure at the interface (e.g., fixed connection, hinge), a second coordinate transformation (independent coordinate transformation) is performed to eliminate non-independent modal coordinates, resulting in a set of independent generalized coordinates describing the motion of the entire structure, composed of the independent modal coordinates of each substructure. This leads to the derivation of the aircraft dynamic equations with aileron clearance nonlinearity in the form of independent modal coordinates for the entire aircraft structure, thus establishing the aircraft dynamic model with aileron clearance nonlinearity.

[0064]

[0065] In the formula, For the quality matrix, Here is the stiffness matrix. It is a non-linear vector. For independent generalized coordinates.

[0066] Step 2: In the aircraft dynamics model with aileron clearance nonlinearity established in Step 1, the additional inertial force caused by the aileron deflection is introduced to obtain the motion equations of the aircraft servo-elastic system with aileron clearance nonlinearity in independent generalized coordinates.

[0067]

[0068] in, The additional inertial force at the structural nodes caused by the deflection motion of the longitudinal rudder surface. For the aileron deflection, The cross mass array, and the aileron deflection mode shape Related, can be approximated as

[0069]

[0070] In the formula, For the rudder axis heading coordinates, Here are the heading coordinates of each node; define the direction along the airflow as... The axis is positive, and the trailing edge of the aileron deflects downwards in the positive direction.

[0071] Step 3: Construct a nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity, which includes the following processes:

[0072] Sub-step 3.1: Establish the response model of the sensor installation location in physical coordinates and obtain the sensor output equation. In the formula, The state vector includes modal coordinates, control surface deflection states, and flight control system state variables. The modal matrix for the sensor mounting location.

[0073] Sub-step 3.2: First, based on the servo motor transfer function, establish the corresponding servo motor dynamic state-space equations.

[0074] ,in, The state vector of the manipulation surface. The control surface deflection command is fed back to the servo motor from the flight control system. and These are the system matrix and control matrix of the servo motor, respectively.

[0075] Then, combining the sensor output equation from step 3.1 and the motion equations of the open-loop servo-elastic system of the aircraft in independent generalized coordinates with aileron clearance nonlinearity established in step 2, the state-space equation form of the open-loop nonlinear servo-elastic motion equation is obtained. ,in and These are the system matrix and control matrix of the servo-elastic system, respectively. It is a non-linear vector.

[0076] Sub-step 3.3: Establish the flight control system model. The state-space equations of the flight control system are as follows: ,

[0077] And set the servo input. With the output of the flight control system The gain feedback relationship between them is .

[0078] In the above formula, This is the state vector of the flight control system. The sensor output vector, , , and These are the system matrix, control matrix, output matrix, and direct transfer matrix of the flight control system. This is the gain matrix.

[0079] Sub-step 3.4: Based on the results of steps 3.1-3.3, obtain the nonlinear servo-elastic closed-loop system equations of the aircraft with aileron clearance nonlinearity, and construct the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity. Figure 3 This is a block diagram of a nonlinear servo-elastic closed-loop system for an aircraft with aileron clearance nonlinearity.

[0080] Step 4: The nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity established in Step 3 is modified to obtain the modified nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity. The specific modification process is as follows:

[0081] Sub-step 4.1: Conduct ground vibration tests on the aircraft using the LMS experimental modal analysis system to obtain modal test results, including modal frequencies and mode shapes. Based on these modal test results, modify the material properties or structural geometric parameters of the aircraft structural dynamics finite element model constructed in step 1 to alter the stiffness and mass characteristics of the dynamics finite element model, ensuring that the modal frequencies and mode shapes are consistent with the subsequent test model.

[0082] Sub-step 4.2: Based on the servo drive control principle, including the servo drive control method, test system, and excitation signal settings, perform servo-rudder deflection angle frequency response test, collect aileron deflection signals and corresponding spectrum curves, and fit the system's transfer characteristics through a multi-order transfer function to obtain the amplitude frequency and phase frequency transfer characteristic curves of the nonlinear servo elastic closed-loop system of the aircraft containing aileron clearance nonlinearity. Correct the servo transfer function in step 3.2 using the obtained transfer characteristic curve functions.

[0083] (1) For the servo drive control method, the control signal of the digital servo is a 50Hz PWM square wave, where the high-level duration corresponds to the servo deflection angle, and the high-level time and the servo deflection angle are linearly related. A high-level duration of 0.5ms corresponds to a servo deflection of 0°, and a high-level duration of 2.5ms corresponds to a servo deflection of 180°, such as Figure 4 The diagram shows the relationship between the duration of the PWM high level and the angle.

[0084] (2) In the experiment, the board of the Quanser hardware-in-the-loop simulation system was selected as the signal generation and acquisition device, and the following were designed: Figure 5The experiment demonstrates the frequency response characteristics of the structure shown. First, a signal modulation module is built in the simulation system to modulate the swept frequency signal into the PWM wave duty cycle required by the PWM generator in the Quanser board, and then sends it to the servo motor. Upon receiving the PWM wave, the servo motor executes a deflection command, driving the aileron to deflect and exciting structural vibration. A laser sensor is connected to the ADC sampling module of the Quanser board to collect the displacement response data of the structure and save the collected data in the simulation system. Based on the response data and the swept frequency signal, the frequency response function can be obtained. Taking a 5° swept frequency of the servo motor as an example... Figure 6 The spectrum analysis diagram of the aircraft wing servo motor at a 5° frequency sweep is given.

[0085] (3) The transfer characteristics of the system are fitted using a multi-order transfer function to obtain, for example: Figure 7 The system amplitude and phase frequency fitting results shown have the following transfer function fitting results:

[0086]

[0087] Step 4.3: Using the "load-displacement curve" clearance measurement method, the aileron is loaded and unloaded in both directions, and its displacement value under the corresponding load is measured. The load-displacement curve of the aileron is plotted, and the clearance displacement under zero load is obtained from the "load-displacement curve". Due to the existence of factors such as friction and damping, the loading and unloading curves cannot be completely coincident. The control surface deflection stiffness is corrected according to the clearance size to accurately characterize the nonlinear stiffness characteristics of the control surface clearance. Its load-displacement curve is as follows: Figure 8 As shown.

[0088] Step 5: Perform simulation analysis on the clearance nonlinearity of the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity, and obtain the simulation analysis results (system response spectrum).

[0089] An open-loop frequency response test was conducted on the modified nonlinear servo-elastic closed-loop system model of the aircraft, which includes aileron clearance nonlinearity, to obtain the test results of servo-elastic characteristics. The simulation analysis results of the servo-elastic characteristics were then compared with the test results. If the simulation analysis results matched the test results, it indicated that the modified nonlinear servo-elastic closed-loop system model of the aircraft, which includes aileron clearance nonlinearity, met the stability requirements. Otherwise, the process returned to step 4 to re-correct the model. The correction parameters included the aircraft structural dynamics characteristics, servo transmission characteristics, clearance nonlinearity parameters, and flight control system parameters. The specific process is as follows:

[0090] By conducting open-loop frequency response tests, Bode and Nyquist plots of the corrected nonlinear servo-elastic closed-loop system of the aircraft, including aileron clearance nonlinearity, can be obtained. From the Bode and Nyquist plots, the system's amplitude-frequency response, phase-frequency response curves, and open-loop stability margin can be derived. Taking a 5° servo sweep frequency test as an example... Figure 9 Bode plots are given for the open-loop servo-elastic system of an aircraft with aileron clearance nonlinearity. Figure 10 Provide the corresponding Nyquist plot of the system. From Figure 9 and Figure 10 As can be seen from the Bode plot and Nyquist plot, the amplitude-phase frequency response curve does not encircle the point (-1, j0), indicating that the constructed servo elastic system is stable. This is common knowledge in the field and will not be explained in detail here.

[0091] To further verify the closed-loop stability and closed-loop stability margin of the aircraft servo-elastic dynamics model established and modified by the method of the present invention, this embodiment also includes a closed-loop servo-elastic test, including three channels: pitch channel, roll channel, and yaw channel. Figure 11 A schematic diagram of the closed-loop coupling test is given. Pulse excitation is applied to the aircraft structure, and the time history of the airframe acceleration signal and the rudder command output signal is recorded under different gain amplification factors in the control channel. The closed-loop stability of the system can be obtained from the attenuation of the output signal over time.

[0092] Taking the pitch channel closed-loop servo elastic test as an example, Figure 12 The time-domain response signal of the aircraft pitch acceleration when the pitch channel feedback gain K=225 is given. Figure 12 It can be seen that the pitch direction shows a significant attenuation of the servo deflection signal (i.e., the angular velocity response signal) at feedback gain K=225, indicating that the closed-loop servo elastic system is stable. Based on this, the feedback gain is gradually increased, and the servo deflection signal is observed. Figure 13 The time-domain response signal of pitch acceleration with pitch channel feedback gain K=450 is given. Figure 13 The results show that the time history of the servo deflection signal (i.e., the angular velocity response signal) exhibits constant amplitude oscillation, indicating that the system is in a critical state.

[0093] Figure 14 A comparison is presented between the experimental results of the open-loop frequency response test and the simulated response spectrum of the system after correction using the servo elastic dynamics model of this invention. Figure 14 It can be seen that the simulation analysis results of the corrected gap nonlinear aircraft servo elastic characteristics are in good agreement with the experimental results, indicating that the model correction method of the present invention has achieved the goal of improving the accuracy of servo elastic analysis. The aircraft servo elastic dynamics model established and corrected by the method of the present invention can reflect the physical characteristics of real aircraft, is suitable for engineering applications, and provides a reliable basis for the aerodynamic servo elasticity evaluation of aircraft.

[0094] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any person skilled in the art can easily conceive of various equivalent modifications or substitutions within the scope of the technology disclosed in the present invention, and such modifications or substitutions should all be covered within the scope of protection of the present invention.

Claims

1. A method for correcting an aircraft servo-elastic dynamics model considering control surface clearance, characterized in that, Includes the following steps: Step 1: Divide the aircraft structure containing aileron clearance nonlinearity into several substructures, and use several nodes of the nonlinear connection region on the aircraft structure as interfaces to discretize each substructure to obtain the corresponding branch modes of each substructure; based on the branch modes of each substructure, establish an aircraft dynamics model containing aileron clearance nonlinearity. Step 2: In the nonlinear aircraft dynamics model with aileron clearance established in Step 1, the additional inertial force caused by the aileron deflection is introduced to obtain the motion equations of the open-loop servo-elastic system of the aircraft with aileron clearance nonlinearity in independent generalized coordinates: In the formula: This is the quality matrix; Here is the stiffness matrix; For independent generalized coordinates; It is a non-linear vector; The additional inertial force at the structural nodes caused by the longitudinal rudder surface deflection motion degree of freedom, and , For aileron deflection; The cross mass array, and the aileron deflection mode shape Regarding this, the aileron deflection mode shape is approximately expressed as: ,in For the rudder axis heading coordinates, Here are the heading coordinates of each node; define the direction along the airflow as... The axis is positive, and the trailing edge of the aileron is deflected downwards to be positive; Step 3: Based on the sensor output equations, the state-space equations of the open-loop nonlinear servo-elastic motion equations, and the state-space equations of the flight control system, establish the nonlinear servo-elastic closed-loop system equations of the aircraft containing aileron clearance nonlinearity, and construct the nonlinear servo-elastic closed-loop system model of the aircraft containing aileron clearance nonlinearity; including: Step 3.1: Establish the response model of the sensor installation location in the physical coordinate system and obtain the sensor output equation. In the formula, The state vector includes modal coordinates, control surface deflection states, and flight control system state variables. The modal matrix for the sensor mounting location; Step 3.2: First, establish the corresponding servo motor dynamic state-space equations based on the servo motor transfer function. ,in, The state vector of the manipulation surface. The control surface deflection command is fed back to the servo motor from the flight control system. and These are the system matrix and control matrix of the servo motor, respectively; Then, combining the sensor output equation from step 3.1 and the motion equations of the open-loop servo-elastic system of the aircraft in independent generalized coordinates with aileron clearance nonlinearity obtained in step 2, the state-space equation form of the open-loop nonlinear servo-elastic motion equation is obtained. , and These are the system matrix and control matrix of the servo-elastic system, respectively. It is a non-linear vector; Step 3.3: Establish the flight control system model. The state-space equation of the flight control system is as follows: And set the gain feedback relationship between the servo input and the flight control system output as follows: ; in, This is the state vector of the flight control system. The sensor output vector, , , and These are the system matrix, control matrix, output matrix, and direct transfer matrix of the flight control system. This is the gain matrix; Step 3.4: Based on the results of Steps 3.1-3.3, obtain the nonlinear servo-elastic closed-loop system equations of the aircraft with aileron clearance nonlinearity, and construct the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity; Step 4: Correct the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity established in Step 3 to obtain the corrected nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity; wherein, the correction parameters include the aircraft structural dynamic characteristics, servo transmission characteristics, clearance nonlinearity parameters and flight control system parameters. Step 5: Perform servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity after modification, and conduct servo-elasticity test and simulation analysis. Obtain the test results and simulation analysis results of the system response spectrum. Compare the simulation analysis results with the test results. If the simulation analysis results are consistent with the test results, it indicates that the modified nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity meets the stability requirements. Otherwise, return to step 4 to modify the model again.

2. The method for correcting an aircraft servo-elastic dynamics model considering control surface clearance as described in claim 1, characterized in that, Step 4, the process of correcting the nonlinear servo-elastic closed-loop system model of the aircraft with aileron clearance nonlinearity established in step 3, includes: Step 4.1: Conduct ground vibration modal tests on the aircraft, obtain modal test results, including information on the modal frequencies and mode shapes of each order; and modify the material properties or structural geometric parameters of the aircraft dynamic model with aileron clearance nonlinearity constructed in Step 1 based on the modal test results, so as to change the stiffness and mass characteristics of the aircraft dynamic model and ensure that the modal frequencies and mode shapes are consistent with the test model. Step 4.2: Perform servo-rudder deflection frequency response test to obtain aileron deflection signal and corresponding spectrum curve. Fit the amplitude frequency and phase frequency transfer characteristic curve function of the nonlinear servo elastic closed-loop system of the aircraft containing aileron clearance nonlinearity. Correct the servo transfer function in step 3.2 using the transfer characteristic curve function. Step 4.3: The aileron is loaded and unloaded in both directions using the "load-displacement curve" gap measurement method to obtain the aileron displacement under the corresponding load and the aileron loading-displacement curve. The gap size under zero load is obtained based on the loading-displacement curve. The control surface deflection stiffness is corrected based on the gap size to accurately characterize the nonlinear stiffness of the control surface gap.

3. The method for correcting an aircraft servo-elastic dynamics model considering control surface clearance as described in claim 2, characterized in that, In step 4.2, the specific process of performing the servo-rudder deflection frequency response test to obtain the aileron deflection signal and the corresponding spectrum curve is as follows: Step 4.2.1: Build a test system, including the Quanser hardware-in-the-loop simulation system, which has a signal PWM generator, a modulation and demodulation module, and an ADC sampling module; A laser sensor is installed in the nonlinear servo-elastic closed-loop system model of the aircraft containing aileron clearance nonlinearity. The laser sensor is connected to the ADC sampling module to collect aileron displacement response data. Step 4.2.2: The Quanser hardware-in-the-loop simulation system modulates and demodulates the input sweep frequency signal and then inputs it to the PWM generator. The demodulated signal satisfies the PWM wave duty cycle required by the PWM generator. The PWM generator outputs a PWM square wave control signal to the servo motor according to the PWM wave duty cycle; wherein the high-level duration of the PWM square wave control signal corresponds to the servo motor deflection angle, and the high-level duration is linearly related to the servo motor deflection angle. Step 4.2.3: The servo executes the deflection command according to the received PWM square wave control signal, driving the aileron to deflect and exciting the aileron to vibrate; Step 4.2.4: The laser sensor collects the aileron displacement response data and transmits it to the AD sampling module of the Quanser hardware-in-the-loop simulation system; Step 4.2.5: Obtain the aileron frequency response function based on the collected displacement response data and the frequency sweep signal.

4. The method for correcting an aircraft servo-elastic dynamics model considering control surface clearance as described in claim 1, characterized in that, In step 5, the servo elasticity test includes an open-loop frequency response test. Based on the open-loop frequency response test, the Bode plot and Nyquist plot of the system are obtained to obtain the open-loop amplitude-phase frequency characteristic curve and the open-loop stability margin of the system.

5. The method for correcting an aircraft servo-elastic dynamics model considering control surface clearance according to claim 4, characterized in that, Step 5 also includes a closed-loop servo elasticity test; the closed-loop servo elasticity test is used to obtain the closed-loop stability margin of the system.

6. The method for correcting an aircraft servo-elastic dynamics model considering control surface clearance as described in claim 1, characterized in that, Step 1 is implemented based on the Craig-Bampton fixed interface modal synthesis method.

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