Aircraft slat fatigue test follow-up angle error analysis and optimization method

Abstract

The analysis and optimization method for the follow-up angle error of the aircraft slat fatigue test comprises the following steps: carrying out the slat fatigue test by adopting the angle error between the driving angle and the frame inclination angle, adjusting parameters of the slat fatigue test through the kinematic model, carrying out the initial slat fatigue test, and carrying out the final slat fatigue test. The method comprises the following steps: acquiring loading control data of an aircraft slat, acquiring a plurality of flight spectrums in an aircraft slat fatigue test, selecting a flight spectrum with the highest occurrence frequency from the plurality of flight spectrums as a to-be-analyzed flight spectrum, transforming clamping positions of all aircraft slats according to the loading control data to obtain an error transformation value of a follow-up angle, and optimizing the error transformation value of the follow-up angle to obtain a to-be-analyzed flight spectrum. According to the method, the test efficiency can be improved.

Description

Technical Field

[0001] This invention belongs to the technical field of full-scale fatigue testing of aircraft slats, and particularly relates to a method for analyzing and optimizing the follow-up angle error in aircraft slat fatigue testing. Background Technology

[0002] In full-scale fatigue tests of aircraft slats, the wing surface deflection motion is usually accomplished by an AC servo motor drive system, while the wing surface load is applied by the control system through a swing-arm type servo loading system. The angle error between the two systems is the most important criterion for the accuracy control of the servo loading system, and it is also an important safety protection measure. Therefore, the optimization of the angle error is of great value.

[0003] Currently, the follow-up loading system is a control method based on displacement driving and angle monitoring. It uses time as the basis for synchronous motion. The angle error depends on the combined influence of factors such as position control, drive system and time difference. Only through repeated iterations and a lot of debugging work can the angle error be kept within the range required by the test accuracy. The debugging process is time-consuming and labor-intensive. It is not targeted or specific for the optimization of angle error and has relatively low efficiency.

[0004] In view of this, the present invention is hereby proposed. Summary of the Invention

[0005] The method for analyzing and optimizing the servo angle error in aircraft slat fatigue testing provided by this invention enables high-precision optimization of the servo loading system's real-time tracking drive system's angle error in slat fatigue testing, thereby improving testing efficiency. The technical solution of this invention has many beneficial effects, as described below: A method for analyzing and optimizing the follow-up angle error in aircraft slat fatigue testing, wherein the aircraft is equipped with a drive system, a follow-up frame displacement system, a frame tilt sensor, and a slat surface tilt sensor, the analysis and optimization method comprising: Step 1: Obtain the driving angle, calculated angle, frame tilt angle, and slat tilt angle fed back by the drive system, follow-up frame displacement system, frame tilt angle sensor, and slat tilt angle, respectively. Use the angle error between the driving angle and the frame tilt angle to conduct a slat fatigue test. Step 2: The frame tilt angle and drive angle are used to construct a kinematic model of the displacement elongation of the actuator cylinder of the follower frame displacement system and the deflection angle of the aircraft slat surface through the mechanical structure digital model of the follower mechanism. Through the kinematic model Adjust the parameters for the slat fatigue test; Step 3: Conduct the initial slat fatigue test and obtain multiple flight spectra from the aircraft slat fatigue test. Select the flight spectra with the highest frequency from the multiple flight spectra as the flight spectra to be analyzed, and continuously collect the loading control data of the flight spectra to be analyzed at a sampling frequency of 128Hz. Based on the loading control data, change the positioning of all aircraft slats to obtain the error transformation value of the follow-up angle. Step 4: Optimize the error transformation value of the follow-up angle until it meets the parameter requirements of the test.

[0006] Compared with the prior art, the technical solution provided by the present invention has the following beneficial effects: This method analyzes four angle values ​​and explicitly uses angle error as the criterion for servo loading control accuracy. Compared to the previous method based on position error, it is more targeted and representative. Simultaneously, it constructs a kinematic model analysis method to theoretically optimize angle error at the initial design stage, effectively avoiding repeated installation problems, reducing workload, and saving time and costs. Furthermore, it employs experimental data analysis to accurately reflect the main sources of angle error throughout the process, serving as an optimization basis. Priorities are assigned according to the degree of influence: position control accuracy optimization, dual-system start-stop synchronization, and periodic correction of the slat surface initial state. This optimization method can be verified in experiments and can serve as an optimization guide for servo systems following slat deflection motion. It can be extended to more servo loading control experiments, showing broad application prospects. Attached Figure Description

[0007] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0008] Figure 1 This is a schematic diagram of the process of the present invention. Detailed Implementation

[0009] The following specific examples illustrate the implementation of the present invention. Those skilled in the art can easily understand other advantages and effects of the present invention from the content disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present invention, and not all of them. The present invention can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed based on different viewpoints and applications without departing from the spirit of the present invention. It should be noted that, in the absence of conflict, the following embodiments and features in the embodiments can be combined with each other. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0010] It should be noted that various aspects of embodiments within the scope of the appended claims are described below. It will be apparent that the aspects described herein can be embodied in a wide variety of forms, and any particular structure and / or function described herein is merely illustrative. Based on this invention, those skilled in the art will understand that one aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number of aspects set forth herein can be used to implement the device and / or practice the method. Additionally, this device and / or method can be implemented using structures and / or functionalities other than one or more of the aspects set forth herein.

[0011] It should also be noted that the illustrations provided in the following embodiments are only schematic representations of the basic concept of the present invention. The drawings only show the components related to the present invention and are not drawn according to the actual number, shape and size of the components in the actual implementation. In the actual implementation, the form, quantity and proportion of each component can be arbitrarily changed, and the layout of the components may also be more complex.

[0012] Furthermore, specific details are provided in the following description to facilitate a thorough understanding of the examples. However, those skilled in the art will understand that aspects can be practiced without these specific details. To enable those skilled in the art to better understand the invention, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. The terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, features defined as "first" and "second" may explicitly or implicitly include one or more of that feature. In the description of the invention, unless otherwise stated, "a plurality of" means two or more.

[0013] In the experimental field, the slat follower frame is generally used as the follower loading system. The position-controlled actuator is connected to the ground, and the force-controlled actuator is connected to the wing surface. The displacement control mode of the coordinated loading control system drives the slat follower frame to realize the angular rotation during positioning and the holding position. The displacement is converted into a calculated angle, and the tilt sensors on the frame and wing surface are used to collect data. The force control mode realizes the follower application of force load. The slat surface drive system is responsible for slat deflection and outputs the deflection angle to the coordinated loading control system for comparison to determine the follower accuracy, that is, the error value of the follower angle.

[0014] like Figure 1 The method for analyzing and optimizing the follower angle error in the fatigue test of an aircraft slat, as shown, involves an aircraft equipped with a drive system, a follower frame displacement system, a frame tilt sensor, and a slat surface tilt sensor. The test is conducted through a position control system, referred to as "position control." The analysis and optimization method includes... Step 1: Obtain the drive angle, calculated angle, frame tilt angle, and slat tilt angle fed back by the drive system, servo frame displacement system, frame tilt angle sensor, and slat tilt angle sensor, respectively. Conduct slat fatigue tests using the angular error between the drive angle and the frame tilt angle to ensure that the servo error angle is maximized during the test. For example, perform big data simulation and modeling by combining the drive angle, calculated angle, frame tilt angle, and slat tilt angle in pairs. Select the drive angle and frame tilt angle with the smallest angular error value as the basis for the test. This not only improves the accuracy of the test but also reduces the cost, the number of tests, and the test cycle.

[0015] Step 2: Construct a kinematic model of the displacement elongation of the actuator cylinder of the follower frame displacement system and the deflection angle of the aircraft slat surface using the mechanical structure digital model of the follower mechanism. Through kinematic models Adjusting the angular error parameters in the slat fatigue test, specifically... and ,in, r1 is the distance from the connection point between the aircraft actuator and the follower mechanism to the axis of rotation of the follower mechanism, and r2 is the distance from the center point of the aircraft actuator projected onto the axis of rotation of the follower mechanism in the deflection plane. This is the initial angle of the aircraft servo mechanism. h The distance from the center point of the actuator cylinder shaft to the deflection plane is denoted as . L 0 represents the initial length of the actuator cylinder; L The displacement that allows the aircraft to be positioned at its maximum or minimum angle; d 0 represents the distance from the center point of the actuator cylinder projected onto the deflection plane to the connection point between the actuator cylinder and the follower mechanism when the error angle is 0°.

[0016] Because there is a certain discrepancy between practice and theory, theoretical principles are used to verify or adjust the fatigue test parameters of the aircraft slats under position control. Furthermore, the follow-up error angle is optimized theoretically first to reduce the overall cost of the test. Specifically... The input quantities of the kinematic model are calibrated. During the pre-test debugging process, the displacement command and loading curve of the load spectrum are continuously modified. The load spectrum is appropriately interpolated during the wing positioning change process to ensure that the angle and angle error at the positioning point meet the test requirements. For example, based on the angle of the test drive system at the aircraft positioning point, the displacement command of the load spectrum at the target positioning point is modified so that the angle error at each positioning point of the aircraft is less than 0.5° and the load spectrum is obtained. The load spectrum is loaded with a linear curve to ensure that the angle error of the follow-up error angle is reduced from 1.5° to 1° during the transformation process.

[0017] Step 3: Conduct the initial slat fatigue test and acquire multiple flight spectra from the aircraft slat fatigue test. Select the flight spectra with the highest frequency from the multiple flight spectra as the flight spectra to be analyzed, and continuously acquire the loading control data of the flight spectra to be analyzed at a sampling frequency of 128Hz. Based on the loading control data, change the positioning of all aircraft slats to obtain the error transformation value of the follow-up angle. Specifically, The positioning transformation of all aircraft slats includes changes in the positioning state, such as preparation, movement, stopping, or holding. For example, preparation (the handshake interaction between the drive system and the servo control system, lasting 2 seconds), movement (the movement of the drive system and the servo control system from the initial positioning to the target positioning, lasting 21 seconds), stopping (the simultaneous arrival of the drive system and the servo control system at the target positioning, entering the error band, the duration of which is determined by accuracy), and holding (the servo system holding at the target positioning, the process of applying force load). Detailed experimental data analysis is performed on these four processes to obtain statistical information such as the maximum and minimum values ​​of angle error, the trend of change, and the error distribution. A macroscopic evaluation of the angle error is made to clarify which specific factors affect the angle error, namely positioning performance, system synchronization, initial state, etc., and the corresponding experimental parameters are adjusted accordingly.

[0018] Step 4: Optimize the error transformation value of the follow-up angle until it meets the parameter requirements of the experiment. Specifically, S1: Feedforward compensation control parameters (feedforward compensation control parameters are special parameters for aviation tests) were added to the control parameters of the loading system of the experimental position control, thereby optimizing the response time and tracking performance of the position control until the accuracy of the experimental position control was 0.1%. S2: The start-up of the position control drive loading system and the wing surface drive system is synchronized. Generally, the start and stop of the position control system are determined through a handshake interaction. The drive loading system performs a preset number of transformations (e.g., 100 times) on all wing surface positions and obtains the transformation time. The average of multiple transformation times is used as the time synchronization basis for position control, and the time parameters are adjusted to ensure that the drive loading system and the wing surface drive system start and stop synchronously and move at the same time. S3: The slat airfoil undergoes a preset number of fatigue tests in its initial state, and the transformed value of the follow-up angle error is obtained. It is then determined whether the transformed value of the follow-up angle error is greater than 1°. If so, the feedforward compensation control parameters in S1 and / or the time synchronization parameters in S2 are adjusted until the transformed value of the follow-up angle error is no greater than 1°. If not, the transformed value of the follow-up error is considered acceptable. All test parameters are output and stored, and the kinematic model is synchronized. It performs parameter updates and outputs theoretical values, etc.

[0019] Overall, this method, also based on displacement-driven and angle-monitoring control, uses in-depth analysis of the entire fatigue test process and loading data to identify the sources of follow-up angle errors at different motion stages, statistically analyze the influence of each source, and provide specific optimization methods for the main sources of follow-up angle errors. This has significant guiding significance for improving the follow-up accuracy and shortening the test cycle in subsequent tests. While maintaining the original test plan, this method analyzes angle errors from multiple perspectives and provides optimization methods, improving implementation efficiency and achieving higher-precision follow-up loading control. It also fully considers the costs of time, funding, and manpower in the test. Using this method, the follow-up angle error can be continuously optimized, ultimately reducing it to within 0.8° (currently, domestic tests use 1° as the standard, and design specifications require it to be within 1°). In other words, the inherent influencing factors of the test have been optimized, effectively improving the efficiency of the test.

[0020] The product provided by this invention has been described in detail above. Specific examples have been used to illustrate the principles and implementation methods of this invention. The descriptions of the embodiments above are merely for the purpose of helping to understand the core ideas of this invention. It should be noted that those skilled in the art can make various improvements and modifications to the invention without departing from the principles of the invention, and these improvements and modifications also fall within the protection scope of the invention claims.

Claims

1. A method for analyzing and optimizing the follow-up angle error in an aircraft slat fatigue test, wherein the aircraft is equipped with a drive system, a follow-up frame displacement system, a frame tilt sensor, and a slat surface tilt sensor, characterized in that, The analysis and optimization methods include, Step 1: Obtain the driving angle, calculated angle, frame tilt angle, and slat tilt angle fed back by the drive system, follow-up frame displacement system, frame tilt angle sensor, and slat tilt angle, respectively. Use the angle error between the driving angle and the frame tilt angle to conduct a slat fatigue test. Step 2: The frame tilt angle and drive angle are used to construct a kinematic model of the displacement elongation of the actuator cylinder of the follower frame displacement system and the deflection angle of the aircraft slat surface through the mechanical structure digital model of the follower mechanism. Through the kinematic model Adjust the parameters for the slat fatigue test; Step 3: Conduct the initial slat fatigue test and obtain multiple flight spectra from the aircraft slat fatigue test. Select the flight spectra with the highest frequency from the multiple flight spectra as the flight spectra to be analyzed, and continuously collect the loading control data of the flight spectra to be analyzed at a sampling frequency of 128Hz. Based on the loading control data, change the positioning of all aircraft slats to obtain the error transformation value of the follow-up angle. Step 4: Optimize the error transformation value of the follow-up angle until it meets the design specifications of the test.

2. The analysis and optimization method according to claim 1, characterized in that, The driving angle, calculated angle, frame tilt angle, and slat tilt angle are combined in pairs for big data simulation and modeling. The driving angle and frame tilt angle with the largest angle error value are selected as the basis for the experiment.

3. The analysis and optimization method according to claim 1, characterized in that, The kinematic model The expression is, and ,in, r1 is the distance from the connection point between the aircraft actuator and the follower mechanism to the axis of rotation of the follower mechanism, and r2 is the distance from the center point of the aircraft actuator projected onto the axis of rotation of the follower mechanism in the deflection plane. This is the initial angle of the aircraft servo mechanism. h The distance from the center point of the actuator cylinder shaft to the deflection plane. L 0 represents the initial length of the actuator cylinder; L The displacement that allows the aircraft to be positioned at its maximum or minimum angle; d 0 represents the distance from the center point of the actuator cylinder projected onto the deflection plane to the connection point between the actuator cylinder and the follower mechanism when the error angle is 0°.

4. The analysis and optimization method according to claim 3, characterized in that, Through the kinematic model Adjusting the angular error parameters in slat fatigue tests includes: Based on the angle of the test drive system at the aircraft position, the displacement command of the load spectrum at the target position is corrected so that the angle error at each position of the aircraft is less than 0.5° and the load spectrum is obtained. The load spectrum uses linear curve loading to ensure that the angle error of the follow-up error angle is reduced from 1.5° to 1° during the transformation process.

5. The analysis and optimization method according to claim 1, characterized in that, Changing the positions of all aircraft slats includes changing the positions to be ready, moving, stopped, or held.

6. The analysis and optimization method according to claim 1, characterized in that, Step 4 includes, Add feedforward compensation control parameters to the control parameters of the loading system of the experimental position control until the accuracy of the experimental position control is 0.1%; The start-up of the position control drive loading system and the wing surface drive system are synchronized in time. The drive loading system performs a preset number of transformations on all wing surface positions and obtains the transformation time. The average value of multiple transformation times is used as the synchronization basis for position control time. The slat surface is subjected to fatigue testing in its initial state, and the change value of the follow-up angle error is obtained. It is determined whether the change value of the follow-up angle error is greater than 1°. If so, the feedforward compensation control parameters and / or time synchronization parameters are adjusted until the change value of the follow-up angle error is not greater than 1°. If not, all test parameters are output and stored.

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