Aircraft structure control coupling test method, apparatus, and device

Abstract

The embodiment of the application provides a kind of aircraft structure control coupling test method, device and equipment, it is related to aircraft field.In this scheme, control device sends excitation signal generation instruction to the flight control of aircraft, and flight control is based on the instruction and the frequency range obtained in advance to generate excitation signal, and it is output to at least one actuator of aircraft.In this process, control device obtains vibration information of aircraft body structure in the process of test by inertial navigation and vibration acquisition device, and judges whether structure control coupling test passes according to vibration information.By this method, the universality of test has been significantly improved, so that it can be applied to different types and specifications of aircraft.

Description

Technical Field

[0001] This application relates to the field of aircraft, and in particular to a method, apparatus and equipment for testing the structural control coupling of aircraft. Background Technology

[0002] In recent years, with the development of materials, energy, and control technologies, the feasibility of developing electric vertical takeoff and landing (eVTOL) aircraft has been fully demonstrated. This type of aircraft combines the advantages of both fixed-wing and rotary-wing aircraft, possessing broad development potential. However, because the overall structural complexity of this aircraft is far greater than that of conventional rotary-wing and fixed-wing aircraft, structural control coupling tests are typically required before the aircraft is put on the market to address the structural control coupling problem.

[0003] Traditional structural control coupling tests are conducted by directly adding excitation sources to the control surfaces, but this method cannot be implemented on rotorcraft. Therefore, a more universal scheme for structural control coupling tests of aircraft is needed. Summary of the Invention

[0004] This application provides a method, apparatus, and equipment for testing the structural control coupling of aircraft, in order to improve the universality of testing the structural control coupling of aircraft.

[0005] In a first aspect, embodiments of this application provide a method for testing the structural control coupling of an aircraft, including:

[0006] Send an excitation signal generation command to the flight control system of the aircraft, the excitation signal generation command being used to instruct the flight control system to generate an excitation signal to drive at least one actuator of the aircraft;

[0007] The vibration information of the aircraft's airframe structure during the test was acquired by the aircraft's inertial navigation and vibration acquisition equipment.

[0008] The structural control coupling test is determined based on the vibration information.

[0009] In one possible implementation, before sending the excitation signal generation command to the flight control system of the aircraft, the method further includes:

[0010] The flight controller is sent excitation configuration information, which carries the frequency range used to generate the excitation signal.

[0011] In one possible implementation, the method further includes:

[0012] Send test status control commands to the flight controller, the test status control commands being used to instruct whether to conduct an open-loop test or a closed-loop test.

[0013] In one possible implementation, the method further includes:

[0014] After the flight controller generates a stabilization control signal and superimposes it onto the excitation signal, it determines whether to adjust the frequency and / or amplitude of the excitation signal based on the excitation signal and the stabilization control signal.

[0015] In one possible implementation, the method further includes:

[0016] If it is determined that the frequency and / or amplitude of the excitation signal needs to be adjusted, a parameter adjustment command is sent to the flight controller, which instructs the flight controller to adjust the frequency and / or amplitude of the excitation signal.

[0017] In one possible implementation, the method further includes:

[0018] An excitation abort command is sent to the flight controller, the excitation abort command being used to instruct the flight controller to stop generating the excitation signal.

[0019] In one possible implementation, determining whether the structural control coupling test passes based on the vibration information includes:

[0020] Based on the vibration information, the excitation response of the aircraft during the test is obtained;

[0021] Based on the excitation response, determine whether the structural control coupling test is successful.

[0022] In one possible implementation, determining whether the structural control coupling test passes based on the excitation response includes:

[0023] The structural control coupling test is deemed passed when the excitation response is greater than the preset amplitude-frequency margin index.

[0024] The structural control coupling test is deemed to have failed when the excitation response is less than or equal to the amplitude-frequency margin index.

[0025] In one possible implementation, the amplitude margin index is 6dB or 9dB.

[0026] In one possible implementation, acquiring the vibration information of the aircraft's airframe structure during the test, obtained by the aircraft's inertial navigation and vibration acquisition equipment, includes:

[0027] During the test, the first vibration information of the aircraft's airframe structure acquired by the vibration acquisition device was received;

[0028] Receive the second vibration information of the aircraft's airframe structure during the test, which is collected by the inertial navigation system and uploaded by the flight control system;

[0029] The vibration information includes the first vibration information and the second vibration information.

[0030] In one possible implementation, the method further includes:

[0031] A test termination command is sent to the flight controller, the test termination command being used to instruct the flight controller to disconnect from at least one actuator.

[0032] In one possible implementation, the frequency range is determined by the characteristics of the aircraft itself and the characteristics of the control system.

[0033] In one possible implementation, the incentive configuration information further includes an incentive type, which is either a single-point incentive or a frequency sweep incentive.

[0034] If the incentive type is frequency sweep incentive, the incentive configuration information also includes frequency sweep step size and frequency sweep count.

[0035] In one possible implementation, the method further includes:

[0036] After completing the test of the aircraft in any configuration, a configuration switching command is sent to the flight controller. The configuration switching command is used to instruct the flight controller to control the aircraft to switch to other configurations for structural control coupling test.

[0037] In one possible implementation, determining whether the structural control coupling test passes based on the excitation response includes:

[0038] In the open-loop test, if the excitation response under single gain is greater than the preset first amplitude margin index, and the excitation response under double gain is greater than the second amplitude margin index, then the structural control coupling test is deemed to have passed.

[0039] Otherwise, the structural control coupling test is deemed unsuccessful;

[0040] The first amplitude margin index is greater than the second amplitude margin index.

[0041] Secondly, embodiments of this application provide a method for testing the structural control coupling of an aircraft, applied to the flight control system of an aircraft, the method comprising:

[0042] Receive excitation signal generation instructions sent by the control equipment;

[0043] An excitation signal is generated according to the excitation signal generation command and the pre-acquired frequency range, and the excitation signal is output to at least one actuator of the aircraft;

[0044] The vibration information collected by the inertial navigation system of the aircraft is returned to the control device.

[0045] In one possible implementation, before the excitation signal sent by the receiving control device generates an instruction, the method further includes:

[0046] The system receives excitation configuration information sent by the control device, wherein the excitation configuration information carries the frequency range.

[0047] In one possible implementation, outputting the excitation signal to at least one actuator of the aircraft includes:

[0048] The excitation signal is distributed according to a preset distribution matrix to obtain the control quantity for each actuator;

[0049] Each actuator is driven according to the control quantity of each actuator.

[0050] In one possible implementation, the method further includes:

[0051] Receive test status control instructions sent by the control device, the test status control instructions being used to indicate whether to perform an open-loop test or a closed-loop test;

[0052] If the test state control command instructs to perform an open-loop test, then the channel of the stabilization control signal is disconnected;

[0053] If the test state control command instructs to perform a closed-loop test, the channel of the stabilization control signal is activated, and a stabilization control signal is generated and superimposed on the excitation signal.

[0054] In one possible implementation, the method further includes:

[0055] Receive parameter adjustment instructions sent by the control device;

[0056] The frequency and / or amplitude of the excitation signal are adjusted according to the parameter adjustment command.

[0057] In one possible implementation, the method further includes:

[0058] Receive the excitation stop command sent by the control device;

[0059] The generation of the excitation signal is stopped according to the excitation stop instruction.

[0060] In one possible implementation, the method further includes:

[0061] Receive the termination test command sent by the control device;

[0062] According to the termination test command, disconnect the connection between the flight controller and the at least one actuator.

[0063] In one possible implementation, the incentive configuration information further includes an incentive type, which is either a single-point incentive or a frequency sweep incentive.

[0064] If the incentive type is frequency sweep incentive, the incentive configuration information also includes frequency sweep step size and frequency sweep count.

[0065] In one possible implementation, the method further includes:

[0066] Receive configuration switching instructions sent by the control device;

[0067] According to the configuration switching command, the aircraft is controlled to switch to other configurations to conduct structural control coupling tests.

[0068] In one possible implementation, the excitation signal is a single-point signal or a frequency sweep signal when performing an open-loop test;

[0069] During closed-loop testing, the excitation signal is either a pulse signal or a frequency sweep signal.

[0070] In one possible implementation, during the closed-loop test, the notch filter is controlled to adjust the stabilization control signal based on the amplitude and phase information of the excitation signal and the stabilization control signal.

[0071] Thirdly, embodiments of this application provide an aircraft structure control coupling test apparatus, comprising:

[0072] The transmitting module is used to send an excitation signal generation command to the flight control system of the aircraft. The excitation signal generation command is used to instruct the flight control system to generate an excitation signal to drive at least one actuator of the aircraft.

[0073] The acquisition module is used to acquire vibration information of the aircraft's airframe structure during the test, obtained by the aircraft's inertial navigation and vibration acquisition equipment.

[0074] The processing module is used to determine whether the structural control coupling test has passed based on the vibration information.

[0075] Fourthly, embodiments of this application provide an aircraft structure control coupling test apparatus, comprising:

[0076] The receiving module is used to receive excitation signal generation instructions sent by the control equipment;

[0077] The processing module is configured to generate an excitation signal based on the excitation signal generation command and a pre-acquired frequency range, and output the excitation signal to at least one actuator of the aircraft.

[0078] The transmission module is used to return the vibration information collected by the inertial navigation system of the aircraft to the control device.

[0079] Fifthly, embodiments of this application provide a computer device, including: a memory, a processor, and an interaction interface;

[0080] The memory stores computer-executed instructions;

[0081] The processor executes computer execution instructions stored in the memory, causing the computer device to perform the aircraft structure control coupling test method as described in any of the first aspects.

[0082] Sixthly, embodiments of this application provide a flight controller, including: a memory, a processor, and an interaction interface;

[0083] The memory stores computer-executed instructions;

[0084] The processor executes computer execution instructions stored in the memory, causing the computer device to perform the aircraft structure control coupling test method as described in any of the second aspects.

[0085] In a seventh aspect, embodiments of this application provide an experimental aircraft, comprising: an aircraft body, the flight control system described in the sixth aspect, and a propeller selected according to formula (1) or formulas (1) and (2) as follows:

[0086]

[0087] Where ρ is the atmospheric density, N is the number of propeller blades, π is pi, and C L R is the propeller thrust coefficient, R is the propeller radius, L is the propeller occupancy information, and N is the propeller's thrust coefficient. RPM d is the nominal rotational speed of the propeller. RPM f is the maximum value of the propeller differential speed. min The lowest sweep frequency is I, where I is the inertia of the aircraft under test, and p is the lowest sweep frequency. max M is the maximum angular rate that drives the aircraft to operate, M is the weight of the aircraft under test, and α is a preset proportional threshold.

[0088] Eighthly, embodiments of this application provide a computer-readable storage medium storing computer-executable instructions, which, when executed by a processor, are used to implement the aircraft structure control coupling test method as described in any one of the first to second aspects.

[0089] Ninthly, embodiments of this application provide a computer program product, including a computer program that, when executed by a processor, implements the aircraft structure control coupling test method as described in any one of the first to second aspects.

[0090] The structural control coupling test method, apparatus, and equipment provided in this application relate to the field of aircraft. In this scheme, the control device sends an excitation signal generation command to the aircraft's flight controller, precisely triggering the flight controller to generate the excitation signal, thereby ensuring the controllability and consistency of the test. The flight controller then generates the excitation signal based on the command and a pre-acquired frequency range, and outputs it to at least one actuator of the aircraft, thus generating a suitable excitation signal according to a specific frequency range, ensuring the accuracy and effectiveness of the test conditions. During this process, the control device acquires vibration information of the aircraft's airframe structure during the test through inertial navigation and vibration acquisition equipment, enabling real-time monitoring of the aircraft's vibration state and providing accurate data support for subsequent analysis. Finally, the structural control coupling test is judged based on the vibration information, thus quickly and accurately determining the test results and ensuring the reliability and safety of the aircraft structure and control system. This method improves the universality of the test, making it applicable to different types and specifications of aircraft, thereby expanding the application scope of the test method and improving its practicality in different aircraft design and testing. Attached Figure Description

[0091] The accompanying drawings, which are incorporated in and form part of this specification, illustrate embodiments consistent with this application and, together with the description, serve to explain the principles of this application.

[0092] Figure 1 A schematic diagram of a scenario for a flight structure control coupling test method provided in this application;

[0093] Figure 2 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 1 ;

[0094] Figure 3 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 2 ;

[0095] Figure 4A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 3 ;

[0096] Figure 5 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 4 ;

[0097] Figure 6 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 5 ;

[0098] Figure 7 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 6 ;

[0099] Figure 8 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 7 ;

[0100] Figure 9 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 8 ;

[0101] Figure 10 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 9 ;

[0102] Figure 11 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 10 ;

[0103] Figure 12 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 10 one;

[0104] Figure 13 This application provides a modal diagram of a conventional fixed-wing aircraft wing with a single bend.

[0105] Figure 14 This application provides a modal diagram of a one-bend structure of an eVTOL wing;

[0106] Figure 15 A modal diagram of a symmetrical pitch structure of an eVTOL external arm provided in this application;

[0107] Figure 16 This application provides a modal diagram of an antisymmetric pitch structure of an eVTOL external arm;

[0108] Figure 17A modal diagram of a symmetrical pitch structure of an eVTOL internal arm is provided in this application;

[0109] Figure 18 This application provides a modal diagram of an anti-symmetric pitch structure of an eVTOL internal arm;

[0110] Figure 19 Modal diagram of a symmetrical swing structure of an eVTOL inner arm provided in this application;

[0111] Figure 20 Modal diagram of an antisymmetric swing structure of an eVTOL internal arm provided in this application;

[0112] Figure 21 A schematic diagram of the host computer interface for a structural control coupling test of an aircraft provided in this application;

[0113] Figure 22 A schematic diagram of an aircraft structure control coupling test system provided in this application;

[0114] Figure 23 A schematic diagram of an aircraft structure control coupling test system provided in this application;

[0115] Figure 24 A closed-loop test schematic diagram provided in this application;

[0116] Figure 25 A schematic diagram of an open-loop test provided in this application;

[0117] Figure 26 A schematic diagram of the framework of an aircraft structure control coupling test system provided in this application Figure 1 ;

[0118] Figure 27 A schematic diagram of the framework of an aircraft structure control coupling test system provided in this application Figure 2 ;

[0119] Figure 28 This is a schematic diagram of the structure of an embodiment of an aircraft structure control coupling test device provided in this application;

[0120] Figure 29 This is a schematic diagram of the structure of a second embodiment of an aircraft structure control coupling test device provided in this application;

[0121] Figure 30 A schematic diagram of the structure of a third embodiment of an aircraft structure control coupling test device provided in this application;

[0122] Figure 31 A schematic diagram of the structure of a computer device provided in this application;

[0123] Figure 32 This is a schematic diagram of a flight control system provided in this application.

[0124] The accompanying drawings illustrate specific embodiments of this application, which will be described in more detail below. These drawings and descriptions are not intended to limit the scope of the concept in any way, but rather to illustrate the concept of this application to those skilled in the art through reference to particular embodiments. Detailed Implementation

[0125] Exemplary embodiments will now be described in detail, examples of which are illustrated in the accompanying drawings. When the following description relates to the drawings, unless otherwise indicated, the same numbers in different drawings denote the same or similar elements. The embodiments described in the following exemplary embodiments do not represent all embodiments consistent with this application. Rather, they are merely examples of apparatuses and methods consistent with some aspects of this application as detailed in the appended claims.

[0126] First, let me explain the terms used in this application:

[0127] Powered aircraft: refers to aircraft that can achieve vertical take-off and landing and low-speed flight, and can cruise like a fixed-wing aircraft;

[0128] Electric Vertical Takeoff and Landing (eVTOL): An aircraft that uses an electric drive system to achieve vertical takeoff and landing, combining the characteristics of fixed-wing and rotary-wing aircraft, and aims to provide an efficient and environmentally friendly air transportation solution;

[0129] Inertial navigation: a navigation system that uses inertial sensors to measure the linear acceleration and angular velocity of an aircraft, thereby calculating its position, velocity and attitude;

[0130] Actuator: refers to the device or component in the control system that is responsible for executing control commands. In an aircraft, the actuator receives signals from the flight control system to adjust the control surfaces or other movable parts of the aircraft, thereby changing the attitude and trajectory of the aircraft.

[0131] Flight control: is an electronic system on an aircraft used to automatically or manually control the aircraft's flight attitude and trajectory;

[0132] Structural-control coupling test: also known as servo-elasticity test, is a test conducted on the ground using the self-excited structural dynamics of the aircraft with its own actuators. It analyzes and verifies the structural-control coupling characteristics of the aircraft. Together with ground vibration test and aerodynamic servo-elasticity simulation, it serves as an important analytical basis for the aerodynamic servo-elasticity evaluation of the aircraft.

[0133] In recent years, with the rapid advancements in materials science, energy technology, and control technology, the feasibility of eVTOL research and development has been fully demonstrated. Due to its environmentally friendly, efficient, and flexible characteristics, eVTOL aircraft have become an important development direction for future urban air transportation. Before bringing such aircraft to market, the coupling problem between its structure and control system must be solved to ensure the safety and reliability of the aircraft. Therefore, conducting structural-control coupling tests is a crucial step in the research and development process.

[0134] Currently, in conducting structural control coupling tests on aircraft, existing technologies typically rely on multiple independent instruments and devices. A sweep frequency generator or signal generator is used as the excitation source, producing a tunable frequency signal to stimulate the aircraft's airframe structure. Oscilloscopes and data loggers are used to capture and record the dynamic response of the airframe structure. These devices are configured and coordinated to achieve synchronous data acquisition and analysis.

[0135] However, the current methods for applying excitation sources in structural control coupling tests are difficult to implement on rotorcraft, and cannot effectively simulate the unique coupling characteristics of their structure and control systems. Therefore, there is an urgent need to develop a more universal method for structural control coupling tests on aircraft to improve the applicability of the tests.

[0136] To address the aforementioned technical problems, the inventors, during their research on structural control coupling test methods for aircraft, discovered that current test methods are primarily applicable to the excitation source application methods for fixed-wing aircraft, and are difficult to effectively apply to rotorcraft or compound-wing aircraft such as eVTOL aircraft. Furthermore, existing structural control coupling test methods require complex equipment, typically necessitating a separate frequency sweeper or signal generator as the excitation source, and relying on equipment such as oscilloscopes to observe and record test results. Based on this, the inventors considered changing the excitation source application method to improve the test's versatility while simplifying the complexity of the test equipment. Specifically, the control equipment directly interacts with the aircraft's flight controller, sending excitation signal generation commands. The flight controller generates excitation signals based on these commands and a preset frequency range, and transmits them to the aircraft's actuators, thereby eliminating dependence on independent excitation equipment and simplifying the test equipment's complexity. Simultaneously, the control equipment acquires real-time vibration information of the aircraft structure through inertial navigation and vibration acquisition equipment to determine the results of the structural control coupling test. This method not only improves the test's versatility, making it applicable to different types and specifications of aircraft, but also ensures the test's accuracy and reliability.

[0137] Figure 1 This is a schematic diagram illustrating a scenario for an aircraft structure control coupling test method provided in this application. Figure 1 As shown, the scenario of this aircraft structure control coupling test method includes control equipment 101 and flight controller 102. Although Figure 1Only one control device 101 and one flight controller 102 are shown, but it should be understood that there may be two or more control devices 101 and one flight controller 102.

[0138] The control device 101 can be a single terminal device, such as a laptop or desktop computer. Alternatively, the control device 101 can be a combination of multiple terminal devices. Furthermore, the control device 101 can also be a standalone server, a cloud server cluster, or other types of equipment. This application does not specifically limit the specific form and type of the control device 101.

[0139] Flight Controller 102 is an electronic system on an aircraft used to automatically or manually control the aircraft's flight attitude and trajectory.

[0140] During the structural control coupling test of the aircraft, the control device 101 sends an excitation signal generation command to the flight controller 102 of the aircraft. The flight controller 102 generates an excitation signal based on the excitation signal generation command and a pre-acquired frequency range, and outputs the excitation signal to at least one actuator of the aircraft. In this process, the control device 101 acquires the vibration information of the aircraft's airframe structure collected by the aircraft's inertial navigation and vibration acquisition equipment during the test, and determines whether the structural control coupling test is passed based on this vibration information.

[0141] It should be noted that the above scenario is only an example of an application scenario provided by the embodiments of this application. The embodiments of this application do not limit the actual form of the various devices included in the scenario. In the specific application of the solution, it can be set according to actual needs.

[0142] The technical solution of this application and how the technical solution of this application solves the above-mentioned technical problems are described in detail below with specific embodiments. These specific embodiments can be combined with each other, and the same or similar concepts or processes may not be described again in some embodiments. The embodiments of this application will now be described with reference to the accompanying drawings.

[0143] Figure 2 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 1 ,like Figure 2 As shown, the method includes:

[0144] S201: Send excitation signal generation command to the flight control system of the aircraft.

[0145] In this step, in order to overcome the limitation that the excitation source implementation method in the existing aircraft structure control coupling test is only applicable to fixed-wing aircraft, this application proposes a general test method.

[0146] Specifically, during the aircraft structure-control coupling test, the control equipment sends excitation signal generation commands to the aircraft's flight controller. It is worth noting that the control equipment connects to the flight controller via a wired or wireless communication interface to send these excitation signal generation commands. Furthermore, the control equipment is also used to configure at least one of the following: excitation signal generation, termination, cancellation, frequency, and amplitude. The generation of the excitation signal marks the start of the test; the cancellation of the excitation signal marks a temporary pause in the test; and the cancellation of the excitation signal marks the end of the test or the completion of a phase. Frequency configuration of the excitation signal refers to selecting a specific single-point frequency or defining a frequency range for frequency sweep testing. Amplitude configuration of the excitation signal refers to setting a fixed amplitude or amplitude range to simulate excitation conditions of varying intensities.

[0147] The main purpose of the aircraft structure-control coupling test is to obtain the structure-control coupling characteristics of the aircraft by conducting servo-elasticity tests on the ground, and to determine whether the structure-control coupling loop is stable. Simultaneously, the test results can be used to correct the mathematical model in the aerodynamic servo-elasticity calculation, making the aerodynamic servo-elasticity analysis more reliable. Furthermore, the aircraft structure-control coupling test is a verification test, conducted after the ground vibration test, focusing on the excitation conditions near the structural modal frequencies.

[0148] The overall structure and various systems of an aircraft are the objects of structural control coupling tests. These aircraft include, but are not limited to, rotorcraft, compound wing aircraft, or fixed-wing aircraft. It is worth noting that the aircraft can be a single aircraft or a combination of multiple aircraft for collaborative testing or comparative analysis. This application does not specifically limit the specific type, structure, and configuration of the aircraft to accommodate different testing needs and application scenarios. Furthermore, it is important to note that the overall structure refers to the physical structure of the entire aircraft, including the airframe structure and other auxiliary structural parts. The various systems refer to the subsystems within the aircraft that are related to structural performance, including but not limited to flight control, navigation systems, communication systems, and actuation systems.

[0149] It should be noted that the main type of aircraft involved in this application refers primarily to powered lift aircraft.

[0150] Flight controllers are primarily used to receive and process excitation signals from control devices and generate commands. Flight controllers include, but are not limited to, embedded controllers, other computing platforms applicable to aircraft control, etc., and this application does not specifically limit them.

[0151] Excitation signals are input signals applied by the flight control system to the actuators during aircraft structure-control coupling tests. Their purpose is to elicit a dynamic response from the aircraft structure. Excitation signals can be pre-set waveforms, such as sine waves, designed to induce a structural response in the aircraft so that its dynamic behavior can be observed and measured.

[0152] The excitation signal generation command is a command issued by the control equipment to the flight controller of the aircraft, which instructs the flight controller to generate an excitation signal to drive at least one actuator of the aircraft.

[0153] The actuators perform corresponding actions based on the excitation signals generated by the flight control system to influence the structural dynamics of the aircraft, thereby achieving the objective of the structural control coupling test. It is worth noting that different types of aircraft have different actuators, and this application does not impose specific limitations on them.

[0154] In one specific implementation, if the aircraft is a fixed-wing aircraft, at least one actuator includes servos respectively located at the elevator, rudder, and aileron positions of the fixed-wing aircraft, or servos respectively located at the elevator rudder and aileron positions of the fixed-wing aircraft. It is worth noting that the fixed-wing aircraft in which at least one actuator is located at the elevator rudder and aileron positions of the fixed-wing aircraft is a V-tail aircraft. The elevator rudder is a control surface on the V-tail of a V-tail aircraft, combining the functions of the elevator and rudder, and is mainly used to control the pitch and yaw of the aircraft.

[0155] If the aircraft is a rotorcraft, at least one actuator includes an electric motor that drives the propeller.

[0156] If the aircraft is a compound wing aircraft, at least one actuator includes: a servo motor located at the elevator, rudder and aileron positions, and a motor driving the propeller; or a servo motor located at the elevator, rudder and aileron positions, and a motor driving the propeller.

[0157] Ailerons are used to control the roll of an aircraft, that is, its rotation about its longitudinal axis. Ailerons are usually mounted on the wings and can enable the aircraft to roll or maintain balance during flight.

[0158] The location of the actuators directly affects the control accuracy and response speed of an aircraft. Placing the actuators near the control surfaces reduces signal transmission delays and improves control response accuracy. If the actuators are far from the control surfaces, more complex transmission mechanisms are required, which not only increases the weight and complexity of the system but may also lead to response delays, affecting the aircraft's handling performance. Therefore, placing the actuators near the control surfaces ensures more efficient and sensitive control.

[0159] It should be noted that before the test begins, the client should provide the testing unit with the necessary technical support documents, mainly including the aircraft's weight, center of gravity, and inertia data, measurement point coordinate data, vibration test results of the aircraft, and center frequency, width, and notch depth data of the notch filters for each channel. The vibration test results of the aircraft should include the mode shapes and frequencies of each mode below 10 times the attitude control bandwidth.

[0160] S202: Receive excitation signal generation command sent by the control device.

[0161] In this step, the aircraft's flight control system receives and sends excitation signals from the control equipment to generate commands.

[0162] The excitation signal generation command is used to instruct the flight control system to generate an excitation signal for at least one actuator that drives the aircraft.

[0163] Optionally, in one specific implementation, after the flight controller receives the excitation signal generation command sent by the control device, it will perform command parsing and command verification operations to ensure that the flight controller can correctly understand and execute the excitation signal generation command, thereby ensuring the correct operation of the aircraft actuators and the smooth progress of the test process.

[0164] S203: Generate an excitation signal according to the excitation signal generation command and the pre-acquired frequency range, and output the excitation signal to at least one actuator of the aircraft.

[0165] In this step, the flight controller generates an excitation signal based on the excitation signal generation command sent by the control device received in step S202, and generates the excitation signal according to the excitation signal generation command and the pre-acquired frequency range, and outputs the excitation signal to at least one actuator of the aircraft.

[0166] In one possible implementation, the frequency range refers to the interval between the lowest and highest frequencies of the excitation signal applied in the aircraft structure control coupling test, typically determined by a ground vibration test (GVT). By appropriately setting the frequency range in the aircraft structure control coupling test, resonant frequencies can be identified and avoided, thereby ensuring the stability and reliability of the aircraft under various dynamic conditions.

[0167] The flight controller uses a built-in excitation signal generation module to generate excitation signals. This module contains various excitation signal generation algorithms and tools, capable of generating the required excitation signals based on excitation signal generation commands and pre-acquired frequency ranges.

[0168] Specifically, the excitation signal generation module first determines the type of excitation signal, such as a sine wave for continuous periodic excitation or a square wave for discrete pulse excitation. Next, it adjusts the frequency of the excitation signal according to the frequency range to ensure it operates within the effective range, while simultaneously setting the amplitude and duration of the excitation signal according to the excitation signal generation command. After debugging and verification, the generated excitation signal is output to at least one actuator of the aircraft to induce a corresponding physical response.

[0169] S204: Return the vibration information collected by the aircraft's inertial navigation system to the control equipment.

[0170] In this step, after the flight controller outputs the excitation signal to at least one actuator of the aircraft, the at least one actuator will perform corresponding physical actions according to the received excitation signal, thereby generating vibration information.

[0171] It is worth noting that vibration information is used to simulate the dynamic behavior of an aircraft under different flight conditions to help test its response characteristics. Vibration information includes dynamic parameters such as the linear acceleration and angular velocity of actuators, as well as the resulting characteristics such as frequency, amplitude, and phase. This data comprehensively reflects the physical response of the aircraft when subjected to excitation signals, providing an important basis for evaluating and optimizing aircraft performance. It is important to note that because the vibration acquisition equipment and the inertial navigation system (INS) are located at different positions on the aircraft, the acquired vibration information differs; the vibration acquisition equipment provides the local structural response, while the INS provides the global motion characteristics.

[0172] During this process, the aircraft's inertial navigation system will collect these vibration information in real time and transmit them back to the control equipment.

[0173] The inertial navigation system (INS) is installed on the aircraft's airframe structure, including primary and backup INS, and is connected to the flight control system via signal lines or a dedicated control bus to provide real-time vibration information. It is worth noting that the data recording frequency of the primary and backup INS should be no less than five times the control bandwidth. It is also important to note that the airframe structure refers to the fuselage of the aircraft, which is a component of the overall aircraft structure.

[0174] Inertial navigation systems (INS) are used to collect vibration information from the aircraft's structure during testing. Specifically, the INS collects linear acceleration and angular velocity.

[0175] S205: Acquire vibration information of the aircraft's airframe structure during the test, obtained from the aircraft's inertial navigation and vibration acquisition equipment.

[0176] In this step, the control equipment acquires and integrates vibration information from different sources to comprehensively evaluate the dynamic behavior of the aircraft during the test.

[0177] Specifically, based on step S204, the control device receives vibration information collected by the aircraft's inertial navigation system and transmitted by the flight control system.

[0178] Meanwhile, the control equipment also acquires vibration information of the aircraft's airframe structure during the test from the vibration acquisition equipment. It is worth noting that the vibration acquisition equipment can provide more detailed vibration characteristic data, such as frequency, amplitude, and phase, helping to identify the response characteristics of the aircraft structure under different excitation conditions.

[0179] By integrating data from inertial navigation and vibration acquisition equipment, the control equipment is able to comprehensively analyze the dynamic performance of the aircraft during the test.

[0180] In one specific implementation, vibration acquisition devices are mounted on the aircraft's airframe structure, and multiple devices are used. Specifically, these devices are positioned at the aircraft's inertial navigation system, each actuator, and at least one location characterizing structural vibration modes. The at least one location characterizing structural vibration modes includes the wingtips and / or tailtips. Wingtips and tailtips are typical structural components of aircraft, typically experiencing complex vibration modes. Wings and tails are crucial aerodynamic control surfaces, responsible for generating lift and maintaining stability during flight. When the aircraft is subjected to excitation, wingtips and tailtips are susceptible to aerodynamic effects, especially during high-speed flight or aggressive maneuvers; the vibration characteristics of these areas often significantly impact the overall structural stability of the aircraft.

[0181] Choosing the wingtip and / or tailtip as vibration data acquisition points allows for better capture of the aircraft's vibration patterns under different flight conditions.

[0182] It should be noted that the specific installation location of the vibration acquisition equipment is determined according to the requirements of the ground vibration test, but the inertial navigation system (INS) and motors must be equipped with vibration acquisition equipment. It is important to note that the vibration acquisition equipment deployed in the test should be able to acquire the following information: it should reflect the vibration along the rotor thrust axis at each propeller motor; it should reflect the vibration in the deflection direction at each control surface rear axle; it should characterize the main mode shapes within 10 times the control bandwidth frequency; and it should reflect the three-axis acceleration at the main and backup INS. The vibration acquisition equipment is an accelerometer, which acquires vibration information by collecting acceleration data.

[0183] The inertial navigation system (INS) is installed near the aircraft's center of gravity or another suitable location to ensure accurate measurement of the aircraft's linear acceleration and angular velocity. The INS primarily provides global motion characteristic data to aid in the analysis of the aircraft's overall dynamic behavior and trajectory.

[0184] S206: Determine whether the structural control coupling test is passed based on the vibration information.

[0185] In this step, the control equipment determines whether the structural control coupling test is passed based on the vibration information of the aircraft's airframe structure acquired by the aircraft's inertial navigation and vibration acquisition equipment during the test.

[0186] Optionally, in one specific implementation, the control equipment first preprocesses the vibration information provided by the vibration acquisition equipment and inertial navigation system to eliminate noise and interference. Next, it converts the preprocessed vibration information into vibration signals through spectrum analysis and transforms them from the time domain to the frequency domain to identify the frequency response characteristics of the airframe structure. Subsequently, modal analysis is performed to extract the modal parameters of the structure, including modal frequencies, modal damping, and modal shapes, to reflect the dynamic behavior of the aircraft airframe structure under the excitation signal. Finally, these parameters are compared with expected parameters to evaluate whether the dynamic performance of the airframe structure meets the design requirements. If the parameters measured under all test conditions are consistent with the expected parameters and within the allowable error range, the structural control coupling test can be considered successful, indicating that the aircraft's structural design can meet the dynamic requirements under actual flight conditions. Otherwise, adjustments to the structural design or test conditions are needed to ensure the safety and performance of the aircraft.

[0187] It is important to note that the safety of both personnel and the test specimens must be ensured during the experiment. Specifically, the test site should be under area control, and the safety of personnel can be guaranteed through protective barriers such as wire mesh. An emergency braking switch capable of cutting off high-energy rotor power should also be provided to ensure the safety of the test specimens. Furthermore, personnel should be assigned to monitor the power system to determine the energy source during the experiment, and safety officers and emergency medical personnel should be available to handle any potential injuries or fatalities.

[0188] The aircraft structure-control coupling test method provided in this application involves the flight controller generating corresponding excitation signals based on a pre-acquired frequency range and excitation signal generation instructions sent by the control equipment. These signals are then transmitted to the actuators to achieve aircraft control and dynamic simulation, ensuring the accuracy and relevance of the excitation signals and improving test precision. During the test, the control equipment collects vibration information of the aircraft's airframe structure in real time using inertial navigation and vibration acquisition devices. This provides real-time dynamic data, helping to comprehensively understand the aircraft's response characteristics under different conditions. By analyzing this vibration information, the control equipment can determine whether the structure-control coupling test has passed, thereby quickly and accurately evaluating the coupling characteristics of the aircraft structure and control system, ensuring the reliability of the test results. In summary, this method not only improves the accuracy and reliability of the test but also enhances its universality through flexible frequency range adjustment, real-time data acquisition and analysis, and compatibility with various actuators. This allows the method to be widely applied to aircraft testing of various types and scales, providing important technical support for aircraft design optimization and performance verification, and helping to ensure the safety and efficiency of aircraft in actual operation.

[0189] Figure 3 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 2 ,like Figure 3 As shown, in Figure 2 Based on the flowchart of the aircraft structure control coupling test method shown, before sending the excitation signal generation command to the aircraft's flight control system, the method further includes:

[0190] S301: Sends excitation configuration information to the flight controller, which carries the frequency range used to generate the excitation signal.

[0191] In this step, before the control device sends the excitation signal generation command to the flight controller of the aircraft, it needs to send excitation configuration information to the flight controller in advance.

[0192] Specifically, the excitation configuration information carries the frequency range used to generate the excitation signal.

[0193] In one possible implementation, the frequency range is determined by the characteristics of the aircraft itself and the characteristics of the control system.

[0194] The aircraft's inherent characteristics include factors such as its structure, materials, and dimensions, which determine its response characteristics at different frequencies. Control system characteristics involve the aircraft's control algorithms and the dynamic characteristics of its actuators, which affect the system's control capabilities and stability at different frequencies.

[0195] It is worth noting that the aircraft's inherent characteristics are primarily determined based on the results of GVT (Generating Dynamics Test). GVT is used to evaluate the dynamic response characteristics of an aircraft structure at different vibration frequencies. In GVT, the aircraft structure is subjected to a series of controlled vibration excitations to measure parameters such as its natural frequencies, modal shapes, and damping characteristics. These measurements provide detailed information about the dynamic behavior of the aircraft structure and are crucial for determining the frequency range of the excitation signal. By analyzing the GVT results, researchers can identify the aircraft's resonance characteristics and dynamic stability at different frequencies, thus providing an accurate frequency range for subsequent excitation signal generation. This process ensures that the excitation signal can effectively simulate the aircraft's dynamic response under actual flight conditions, providing a solid foundation for the success of the experiment and the accuracy of the data.

[0196] An aircraft's control system includes flight control, navigation, communication, and actuation systems. The characteristics of an aircraft's control system depend on the type, design, and purpose of the aircraft, and are not specifically limited here.

[0197] By combining the characteristics of the aircraft's own structure and the characteristics of its control system to determine the frequency range, it is possible to ensure that the generated excitation signal matches the actual dynamic behavior of the aircraft. This method allows the flight control system to quickly generate an excitation signal that meets the experimental requirements after receiving the excitation signal generation command, thereby effectively simulating the aircraft's response under various dynamic conditions.

[0198] It should be noted that the process of sending excitation configuration information to the flight control system can be achieved in several ways. It can be executed manually by the test personnel triggering the control equipment, or it can be executed automatically according to preset rules or time nodes. This application does not impose specific limitations on either method. The manual operation method allows the test personnel to flexibly send excitation configuration information based on real-time observation and specific test requirements. This flexibility facilitates real-time adjustments and optimizations during the test, ensuring that test conditions adapt promptly to the dynamic response characteristics of the aircraft. The automatic execution method triggers information transmission through preset rules or time nodes. This automated process improves the efficiency and consistency of the test, reduces errors that may be caused by human operation, and ensures accurate execution under complex test conditions. By combining these two operation methods, the test can achieve a balance between flexibility and efficiency, thereby more effectively simulating the aircraft's response under various dynamic conditions, providing a solid foundation for the success of the test and the accuracy of the data.

[0199] S302: Receive excitation configuration information sent by the control device, the excitation configuration information carrying the frequency range.

[0200] In this step, based on step S301, the control device sends excitation configuration information to the flight controller, which then receives the excitation configuration information sent by the control device. The excitation configuration information includes a frequency range.

[0201] Next, the control device generates an excitation signal based on the frequency range and the excitation signal sent by the flight controller. The specific execution process can be found in steps S203 to S206, and will not be described in detail here.

[0202] In one possible implementation, the incentive configuration information also includes an incentive type, which is either a single-point incentive or a frequency sweep incentive.

[0203] If the incentive type is frequency sweep incentive, the incentive configuration information also includes the frequency sweep step size and the number of frequency sweeps.

[0204] Specifically, in addition to the frequency range, the incentive configuration information also includes the incentive type.

[0205] The incentive types are either single-point incentives or frequency-sweep incentives.

[0206] Single-point excitation is a method of applying vibration at a specific frequency to study the dynamic response of an aircraft's structure at that frequency. This type of excitation allows for focused analysis of the aircraft's behavior at known resonant frequencies or other critical frequencies, helping to identify local characteristics and potential resonance problems in the aircraft's structure.

[0207] Frequency sweep excitation is a method of continuously applying vibrations across a range of frequencies to comprehensively evaluate the dynamic characteristics of an aircraft's structure. The excitation configuration information for frequency sweep excitation includes the sweep step size and the number of sweeps; these parameters determine the level of detail in the frequency variations and the repeatability of the test. Frequency sweep excitation can identify the resonant characteristics and dynamic stability of an aircraft across the entire frequency range, providing extensive frequency response data.

[0208] It should be noted that frequency sweep excitation includes step sweep and continuous sweep. Step sweep is a specific implementation of frequency sweep excitation. Step sweep refers to the frequency gradually increasing or decreasing within a preset range in fixed steps, changing only a fixed frequency value each time. The change in step sweep is abrupt, meaning the frequency change is phased, jumping to the next frequency point each time. Continuous sweep refers to the frequency changing smoothly and continuously within a preset frequency range. Unlike step sweep, the frequency change in continuous sweep is smooth, without interruption or jumps, but rather follows a continuous curve.

[0209] Optionally, the excitation configuration information also includes other parameters, such as excitation amplitude and phase, which also affect the characteristics of the excitation signal and the accuracy of the test results. By comprehensively considering these parameters, test personnel can better control the test conditions and ensure that the excitation signal can effectively simulate the dynamic response of the aircraft in actual operation.

[0210] The aircraft structure control coupling test method provided in this embodiment can detect the effectiveness of aircraft dynamic characteristic tests from multiple dimensions by setting different excitation types. Single-point excitation allows testers to analyze the aircraft's response characteristics at a specific frequency, identifying and resolving local resonance problems. Frequency sweep excitation provides a comprehensive assessment of the aircraft's dynamic characteristics across the entire frequency range, helping to identify resonance phenomena and dynamic stability issues at multiple frequencies. By flexibly selecting and combining these excitation types, the test can better adapt to different analytical needs, thereby improving the test's versatility.

[0211] The aircraft structure control coupling test method provided in this application involves the control device sending excitation configuration information carrying a frequency range to the flight controller to assist the flight controller in generating excitation signals suitable for the test requirements. In this way, the flight controller can accurately generate excitation signals according to a pre-set frequency range, thereby ensuring the accuracy and effectiveness of the test.

[0212] Figure 4 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 3 ,like Figure 4 As shown, based on the above embodiments, the method of outputting an excitation signal to at least one actuator of the aircraft includes:

[0213] S401: Distribute the excitation signal according to the preset distribution matrix to obtain the control quantity of each actuator.

[0214] In this step, after the flight controller generates the excitation signal, it distributes the excitation signal according to the preset distribution matrix to determine the specific control quantity of each actuator.

[0215] Specifically, in the aircraft structure control coupling test, the generation and distribution of excitation signals are designed to simulate the real response of the aircraft under various dynamic conditions.

[0216] The allocation matrix is ​​designed to decompose the overall excitation signal into control signals suitable for each actuator. When designing the allocation matrix, the structural characteristics of the aircraft, the location and function of each actuator, and the specific objectives of the experiment must be considered. Each element of the matrix represents a scaling factor, indicating how much of the overall excitation signal should be allocated to a specific actuator.

[0217] After the excitation signal is generated, the flight controller inputs it into a preset allocation matrix. Through matrix operations, the overall excitation signal is decomposed into multiple control signals, each corresponding to an actuator. The strength and phase of these control signals are precisely calculated to ensure that each actuator responds as expected.

[0218] S402: Drive each actuator according to the control quantity of each actuator.

[0219] In this step, according to step S401, after the flight controller obtains the control quantity of each actuator based on the preset allocation matrix, it will further drive each actuator according to the control quantity of each actuator.

[0220] Specifically, after receiving the control input from each actuator, the flight controller generates corresponding control signals and sends these signals to the drive units of each actuator. The task of these drive units is to translate the received control signals into actual physical actions. For example, the control surface actuators adjust their angles according to the control signals to simulate the aircraft's response under specific dynamic conditions. This precise signal transmission and conversion process ensures that each actuator can accurately perform its intended function, thus providing reliable data support for testing.

[0221] During the propulsion process, the aircraft's inertial navigation system and vibration acquisition equipment installed on the aircraft's airframe structure collect vibration information of the aircraft's airframe structure in real time and transmit it to the control equipment. The specific execution process can be referred to in steps S205 to S206, which will not be repeated here.

[0222] Furthermore, to realistically simulate the flight environment of an aircraft in the air and prevent the ground from affecting the aircraft's operation, in one embodiment of this application, before driving each actuator according to the control quantity of each actuator, the method further includes: placing the aircraft to be tested on a support device, wherein the support device includes an elastic suspension device and a support frame, which can prevent the aircraft from directly rubbing against the ground, thus avoiding inaccurate test results. It is worth noting that the support equipment and support method should be reasonably selected according to the weight of the aircraft and the lowest resonant frequency of the aircraft. Specifically, to avoid resonance, in one embodiment of this application, the frequency of the support device is lower than the preset first-order frequency threshold of the first-order structural modal frequency of the aircraft. Preferably, the first-order frequency threshold is 1 / 3. Specifically, the first-order structural modal frequency of an aircraft refers to the first-order natural vibration frequency generated by the aircraft when subjected to external force or vibration excitation, that is, the free vibration frequency value of the aircraft as a whole around a certain axis or point under the condition of no external constraints.

[0223] The frequency threshold is a preset boundary value, usually a reference frequency set during the design process to take into account the structural strength, stability and safety of the aircraft.

[0224] If the frequency of the support device is higher than or equal to the first-order structural mode frequency of the aircraft, the vibration of the support device may resonate with the natural vibration of the aircraft structure, causing undesirable structural vibrations in the aircraft, thereby affecting the stability and control accuracy of the aircraft. However, when the frequency of the support device is lower than the first-order structural mode frequency of the aircraft, the vibration frequency of the support device will not resonate with the natural frequency of the aircraft, thus avoiding severe vibrations caused by resonance, which helps to improve the stability of the aircraft structure and the accuracy of test data.

[0225] Therefore, by designing the frequency of the support device to be lower than the first-order structural mode frequency of the aircraft, it can be ensured that the support device will not interfere with the natural vibration mode of the aircraft structure during the test, thereby ensuring the stability of the aircraft and the accurate testing of its control coupling characteristics during the test.

[0226] The flexible suspension system consists of air springs or elastic cables. The support frame is a specially designed support frame or the landing gear of the aircraft. By combining the flexible suspension system and the support frame, open-loop and closed-loop tests under different flight configurations can be effectively supported.

[0227] Specifically, the elastic suspension device is used to suspend the aircraft, facilitating open-loop and closed-loop tests of takeoff and landing configurations, transition configurations, and cruise configurations. This helps to accurately simulate the dynamic behavior and response characteristics of the aircraft under different flight conditions. The support frame, on the other hand, is specifically used to support the aircraft in takeoff and landing configuration tests, ensuring that the aircraft is stably fixed during the test and avoiding the impact of external disturbances on the data.

[0228] Air springs, with their adjustable stiffness and damping characteristics, can accurately simulate the dynamic characteristics of aircraft under different flight conditions. They are suitable for test scenarios requiring high-precision control, especially when simulating complex vibration modes and dynamic loads. The use of air springs can effectively reduce ground interference and provide a more realistic flight environment simulation. It is worth noting that the support equipment for air springs mainly consists of the air spring system, lifting supports, and transition connectors.

[0229] Elastic ropes offer a simpler and more flexible suspension method, suitable for tests with less stringent vibration characteristic requirements. They can be quickly deployed and adjusted, and are suitable for preliminary testing and rapid response assessment under various test conditions. In one specific implementation, the elastic rope is a rubber rope, and the support equipment for suspending the rubber rope mainly consists of the rubber rope and transition connectors.

[0230] When selecting a flexible suspension device, the specific test requirements and the characteristics of the aircraft must be considered. For example, air springs are a more suitable choice for tests requiring high-precision dynamic response measurements, while elastic ropes are more suitable for scenarios requiring rapid deployment and flexible adjustment. By rationally selecting and configuring the suspension device, the safety and effectiveness of the test can be ensured.

[0231] Optionally, the support frame can be a specially designed support frame or the landing gear of an aircraft.

[0232] Specially designed support frames typically offer greater stability and adjustability, making them suitable for test scenarios requiring precise control of aircraft attitude and position. These frames provide additional structural support during testing, simulating different load conditions and external environmental influences to ensure the accuracy and consistency of test data. They are suitable for complex open-loop and closed-loop tests, especially when simulating extreme environments or special load conditions.

[0233] Using aircraft landing gear as a support frame offers a more economical and convenient option. Landing gear supports are suitable for tests with less stringent requirements for support stability, can be quickly set up and adjusted, and are suitable for routine takeoff and landing configuration testing and preliminary control system evaluation.

[0234] When selecting a support frame, it is necessary to comprehensively consider the accuracy requirements of the test, the structural characteristics of the aircraft, and the conditions of the test environment. For tests requiring high precision and high stability, a specially designed support frame is a more suitable choice, while for routine testing and rapid deployment needs, using landing gear is more practical. By appropriately selecting the type of support frame, the effectiveness and safety of the test can be ensured.

[0235] Support equipment is configured according to different support methods. When supported by air springs, the support equipment mainly consists of an air spring system, lifting supports, and transition connectors. When suspended by rubber ropes, the support equipment mainly consists of rubber ropes and transition connectors. The support stiffness should ensure that the translational and rotational modal frequencies are less than 1 / 3 of the first-order structural modal frequencies of the aircraft.

[0236] It is worth noting that if a large horizontal force excitation is required during the test, the safety of the aircraft under the excitation should be carefully considered to avoid landing gear support instability or rubber cable suspension system failure. If necessary, horizontal limiting ropes or rubber cables should be added to connect the aircraft body to improve the horizontal stability of the test setup. The added horizontal limiting ropes or rubber cables should not have a significant impact on the aircraft's structural modes.

[0237] During the test, the test equipment should also include control surface elastic fixtures and / or counterweights to increase the excitation force of the control surface during frequency sweep; the control surface elastic fixtures are set on the control surface of the aircraft; the counterweights are set on the rear rafter of the control surface of the aircraft.

[0238] Specifically, the rotational inertia and actuation speed of aircraft control surfaces are limited, and due to nonlinear elements such as actuation clearance, it is difficult to achieve an amplitude-independent excitation effect compared to a rotor. Therefore, it is considered to attach counterweight strips to the rear rafters of the control surfaces to increase the excitation force by increasing the rotational inertia. After adding counterweights, the structural strength of the aircraft should be reassessed, and sufficient structural margin should be reserved (not exceeding 1 / 4 of the structural load limit). The 180° phase difference between the inertial loading of the control surfaces and the aerodynamic loading during flight must be considered. If strict phase requirements are required, elastic loading fixtures for the control surfaces should be used to achieve excitation loading that is approximately in phase with the aerodynamic force.

[0239] The combination of elastic tooling and counterweights on control surfaces can enhance the vibration response of control surfaces in frequency sweep tests, helping to more clearly analyze the coupling effect between the aircraft structure and control system. The elastic tooling improves the stiffness and deformation capacity of the control surfaces, while the counterweights enhance the sensitivity and accuracy of the tests by altering the mass distribution and aerodynamic interactions. The combination of these two elements effectively simulates the dynamic environment during flight, avoids resonance phenomena, and improves the reliability of test data, providing crucial support for evaluating the dynamic behavior of aircraft under different frequency excitations.

[0240] The aircraft structure control coupling test method provided in this application significantly improves the accuracy and reliability of the test by precisely determining the control quantities of the actuators. Using an allocation matrix, this method ensures the accurate allocation of excitation signals, making the actuator response meet expectations.

[0241] Figure 5 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 4 ,like Figure 5 As shown, based on the above embodiments, the method further includes:

[0242] S501: Sends test status control commands to the flight controller. These commands are used to instruct whether to conduct an open-loop or closed-loop test.

[0243] In this step, for each configuration, tests under different load conditions should be conducted as needed, based on the assessment of the aircraft's load conditions. Tests for each configuration should include, but are not limited to, open-loop and closed-loop tests. Therefore, the control equipment needs to send test status control commands to the flight controller to indicate whether an open-loop or closed-loop test is currently being performed.

[0244] Specifically, in aircraft structure-control coupling tests, for aircraft with different configurations, appropriate load states need to be selected and designed for testing based on load assessments under various flight conditions to ensure stable operation of the aircraft in different flight environments. A load condition refers to the external load environment experienced by an aircraft at different flight stages or states. For example, an aircraft encounters different types and magnitudes of loads during takeoff, cruise, maneuvering, stall, and other flight conditions, which affect the aircraft's dynamic response and structural deformation. A load state refers to the specific magnitude, distribution, and temporal variation of the loads experienced by the aircraft under a specific load condition. The test needs to simulate and test the impact of various load states on the structure and control system according to different aircraft configurations.

[0245] The tests should include tests under various load conditions, including but not limited to open-loop and closed-loop tests.

[0246] Specifically, the open-loop test, also known as the open-loop stability margin reserve test, generally uses a swept frequency signal or a single-point signal to excite each channel and perform frequency domain analysis. During the test, the swept frequency amplitude is gradually changed to estimate the stability margin reserve of each channel of the aircraft.

[0247] In open-loop testing, the flight controller operates the actuators according to control commands issued by the control equipment, and does not automatically adjust based on the real-time status of the aircraft. The purpose of designing open-loop tests is to conduct preliminary studies on the structural response of the aircraft, such as wing deformation and vibration modes, and to verify whether the structural design of the aircraft meets the requirements.

[0248] Closed-loop testing, also known as closed-loop stability margin verification, generally involves using pulse or frequency sweep excitation to stimulate each channel and performing time-domain analysis. During the test, the frequency sweep amplitude is gradually changed to verify the stability margin reserve of each channel of the aircraft.

[0249] In closed-loop testing, the flight control system operates the actuators according to control commands issued by the control equipment, monitors the aircraft's status in real time, and automatically adjusts based on feedback information. The purpose of designing closed-loop testing is to verify the stability and accuracy of the flight control system, ensuring that the aircraft can dynamically adjust according to real-time conditions during actual flight, thereby guaranteeing the aircraft's safety, performance, and responsiveness.

[0250] It is worth noting that in the testing process for each aircraft configuration, open-loop testing is conducted first, followed by closed-loop testing. This is because open-loop testing is mainly used to initially verify the structural response and basic performance of the flight control system, while closed-loop testing is conducted based on open-loop testing to evaluate the stability and accuracy of the flight control system under real-time feedback adjustment. This sequence ensures that the flight control system can effectively make dynamic adjustments during actual flight, thereby guaranteeing the safety and performance of the aircraft.

[0251] In one possible implementation, when conducting structural control coupling tests on an aircraft, one can choose to first complete open-loop tests for all configurations, and then conduct closed-loop tests for all configurations. Another approach is to immediately conduct closed-loop tests for a specific configuration after completing the open-loop test, and so on, until all configuration tests are completed.

[0252] It should be noted that the flight controller should be able to send excitation signals to the actuators, and excitation commands can be sent to the flight controller from the host computer. The excitation signals that the flight controller can output should include: step and impact signals of specific amplitude, and sinusoidal sweep signals of specific amplitude and frequency. During the test, the aircraft is subjected to three forms of external force loading: aerodynamic loading, inertial loading, and elastic loading. Aerodynamic loading refers to the force generated by the aircraft driving the power vector actuators (generally rotors). Inertial loading refers to the force generated by the aircraft driving the control surfaces. When the aircraft is in a zero dynamic pressure environment, the periodic command loading of the control surfaces will only generate an inertial force with a 180° lag, failing to produce the desired aerodynamic force. Elastic loading requires the addition of control surface fixtures to connect the control surfaces. These fixtures, without interfering with other structures, allow the control surfaces to obtain additional elastic restoring force when deviating from the zero position. This restoring force is almost in phase with the desired aerodynamic loading force, but has a 180° phase difference with the inertial force.

[0253] It should be noted that the excitation in the experiment is achieved by the aircraft's own actuators, including the rotor and control surfaces. For larger rotorcraft, to ensure experimental safety, the rotor is replaced with a scaled-down smaller rotor. It is necessary to maintain the rotational speed and control bandwidth consistent with the actual operating conditions, while achieving an excitation effect in the frequency domain without generating excessive action force and induced airflow (the total thrust at nominal speed should be less than 1 / 4 of the total weight of the aircraft), in order to ensure the safety of the aircraft and personnel.

[0254] S502: Receives test status control instructions sent by the control equipment. The test status control instructions are used to indicate whether to perform an open-loop test or a closed-loop test.

[0255] In this step, based on step S501, the flight controller receives the test status control command sent by the control equipment. The test status control command is used to indicate whether to conduct an open-loop test or a closed-loop test.

[0256] If the test status control command instructs to perform an open-loop test, then proceed to step S503.

[0257] If the test status control command instructs to perform a closed-loop test, then proceed to step S504.

[0258] S503: If the test status control command indicates that an open-loop test should be performed, then the channel of the stabilization control signal will be disconnected.

[0259] In this step, according to step S502, if the test status control command sent by the control device indicates that an open-loop test should be performed, the channel of the stabilization control signal is disconnected.

[0260] The channel for the stability augmentation control signal refers to the communication path or interface used to transmit the stability augmentation control signal. The stability augmentation control signal is usually generated by the stability augmentation control law in the flight control system, and maintains the stability and handling performance of the aircraft by adjusting parameters such as attitude, speed, and acceleration.

[0261] The stability augmentation control signal is part of the closed-loop control system. It dynamically adjusts the control quantity based on the real-time status feedback of the aircraft to ensure that the aircraft maintains the desired flight state under various flight conditions.

[0262] During open-loop testing, the flight control system does not rely on real-time status feedback for adjustments. Instead, it drives the aircraft's behavior by generating commands through excitation signals sent by the control equipment. In this case, to prevent the stabilization control signal from interfering with the predetermined control commands during the test, the stabilization control signal channel needs to be disconnected. With the stabilization control signal channel disconnected, the aircraft's attitude and state control are no longer adjusted through the stabilization control law, but instead rely entirely on test commands or the control of the test personnel to accurately evaluate the aircraft's performance under specific control conditions.

[0263] S504: If the test state control command indicates that a closed-loop test should be performed, the channel of the stabilization control signal will be turned on, and a stabilization control signal will be generated and superimposed on the excitation signal.

[0264] In this step, according to step S502, if the test status control command sent by the control device indicates that a closed-loop test should be performed, the channel of the stabilization control signal is turned on, and the stabilization control signal is generated and superimposed on the excitation signal.

[0265] Specifically, closed-loop testing occurs in the flight control system of an aircraft, where control signals are dynamically adjusted based on the real-time status feedback of the aircraft. During this process, parameters such as the aircraft's attitude, speed, and acceleration are affected by the stabilization control signals to ensure that the aircraft can operate stably under various flight conditions and maintain the predetermined flight state.

[0266] The stability augmentation control signal is generated by the stability augmentation control law in the flight controller and adjusts the aircraft's state in real time through the flight controller. It is typically used to compensate for environmental disturbances or dynamic changes in the aircraft, ensuring flight stability. When the test state control command instructs a closed-loop test, it is necessary to ensure that the stability augmentation control signal channel is open so that the signal can act on the flight controller, thereby achieving closed-loop control. In one possible implementation, during the closed-loop test, the flight controller receives excitation responses from the control equipment. These responses are the aircraft's dynamic reactions to the input excitation signals, reflecting the system behavior under current flight conditions. The flight controller analyzes these excitation responses and uses a preset algorithm to identify the aircraft's stability state and potential deviations. Specifically, the preset algorithm includes, but is not limited to, any one or any combination of frequency domain analysis, time domain response analysis, and pattern recognition techniques. This application does not impose specific limitations on this; it can be adjusted and determined according to actual needs. Based on in-depth analysis of the excitation responses, the flight controller generates a stability augmentation control signal, a compensation signal used to enhance system stability and correct any trends that may lead to aircraft instability. Specifically, the generation of stabilization control signals involves adjusting the control gain, changing the control law parameters, or introducing additional control inputs to counteract instability factors.

[0267] The generated stabilization control signal is superimposed on the excitation signal in real time, which is achieved through a signal synthesizer to ensure that the stabilization control signal and the excitation signal are precisely matched in time and amplitude.

[0268] It is worth noting that in closed-loop testing, the stability augmentation control signal not only serves as a feedback signal for aircraft stability adjustment but also collaborates with the excitation signal. The excitation signal is used to trigger specific dynamic responses or test scenarios, while the stability augmentation control signal is superimposed on these excitation signals to work synergistically, ensuring that the aircraft can cope with the excitation conditions during the test while maintaining its stability and handling performance.

[0269] Closed-loop testing enables the flight control system to adjust its flight status based on real-time feedback by activating the stability augmentation control signal channel. The superposition of the stability augmentation control signal and the excitation signal helps to accurately evaluate the performance and response of the aircraft under specific test conditions.

[0270] In one specific implementation, a switch 3 is provided on the output path of the stabilization control signal. The switch 3 is located before the excitation signal and the stabilization control signal are superimposed. The switch 3 is used to control the state of the test as either a closed-loop test or an open-loop test, thereby controlling the overall feedback mechanism of the system.

[0271] In open-loop operation, switch 3 is open, and the stabilization control signal is not transmitted to the system's feedback path. In closed-loop testing, switch 3 remains closed, allowing the stabilization control signal to be superimposed on the excitation signal and fed back into the system. Switch 3 is managed by a control device. The control device determines the on / off state of switch 3 based on whether the test is open-loop or closed-loop. This flexible control mechanism allows the system to switch states under different test conditions to meet various testing and analysis needs.

[0272] The structural control coupling test method for aircraft provided in this embodiment can flexibly select open-loop or closed-loop control states according to the actual needs of the test, thereby providing more comprehensive and accurate test results for the aircraft's structure and control system. By combining open-loop and closed-loop tests, this structural control coupling test method not only improves the accuracy of aircraft testing but also effectively reduces test risks, ensuring that the aircraft has good stability and reliability under various operating conditions.

[0273] In one possible implementation, the excitation signal is a single-point signal or a frequency sweep signal when performing an open-loop test;

[0274] During closed-loop testing, the excitation signal is either a pulse signal or a frequency sweep signal.

[0275] Specifically, the choice of excitation signal directly affects the validity and accuracy of the test results when conducting open-loop and closed-loop tests.

[0276] Specifically, in open-loop testing, the excitation signal is either a single-point signal or a frequency-sweep signal. This is because the purpose of open-loop testing is to test the basic dynamic response of the aircraft without any feedback control. A single-point signal is used to excite the aircraft's response at a specific frequency, helping testers analyze the aircraft's natural characteristics, resonance phenomena, and natural frequencies at that frequency. A frequency-sweep signal, on the other hand, covers multiple frequency ranges to test the aircraft's dynamic response at different frequencies. This method can reveal the aircraft's frequency response characteristics and identify potential resonance modes, structural weaknesses, or control problems. Therefore, choosing single-point and frequency-sweep signals as excitation signals helps to understand the aircraft's dynamic characteristics in open-loop conditions in detail, providing data support for further control optimization.

[0277] During closed-loop testing, the excitation signal is either a pulse signal or a swept-frequency signal. This is because the goal of closed-loop testing is to simulate the dynamic response of an aircraft in a real flight environment. A pulse signal is a signal that changes rapidly over a short period of time, characterized by its very short duration and the ability to vary significantly in amplitude within a certain timeframe, thus stimulating the aircraft's frequency response. A swept-frequency signal, on the other hand, covers multiple frequency ranges to test the aircraft's response characteristics at different frequencies.

[0278] In one specific implementation, the pulse signal is a triangular pulse signal. Specifically, a triangular pulse signal is a signal with good time and frequency characteristics, capable of covering multiple frequencies in a short time to stimulate different response modes of the aircraft. By applying a triangular pulse signal, the instantaneous changes that the aircraft may encounter during flight can be simulated, allowing for the evaluation of the flight control system's real-time response capability and stability. Because the triangular pulse signal is symmetrical in signal strength and has a uniform frequency distribution, it allows for a comprehensive evaluation of how the flight control system effectively controls the aircraft under dynamic conditions, ensuring the aircraft remains stable in various changing flight environments.

[0279] It should be noted that the sweep frequency signal and the single-point signal are mutually exclusive. Those skilled in the art can adaptively select one of them as the excitation signal according to the actual experimental needs, and no limitation is made here.

[0280] The aircraft structure control coupling test method provided in this application evaluates the frequency response characteristics of an aircraft under no-feedback conditions by using single-point signals or swept-frequency signals in open-loop tests. In closed-loop tests, pulse signals are used to evaluate the transient response and stability of the aircraft, and swept-frequency signals are used to identify the frequency response characteristics of the aircraft, ensuring the effectiveness and stability of the flight control system at various frequencies, thereby providing comprehensive data support for optimizing flight control design.

[0281] Figure 6 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 5 ,like Figure 6 As shown, based on the above embodiments, the method further includes:

[0282] S601: After the flight controller generates a stability augmentation control signal and superimposes it onto the excitation signal, it determines whether to adjust the frequency and / or amplitude of the excitation signal based on the excitation signal and the stability augmentation control signal.

[0283] In this step, when the aircraft structure-control coupling test is in a closed-loop state, the stabilization control signal generated by the flight controller is superimposed on the excitation signal. The purpose of this operation is to enhance the stability and control accuracy of the aircraft during the test by combining the stabilization signal. However, the superimposed signal may cause the controlled object to make a series of adaptive adjustments at the current frequency. These adjustments may cause brief fluctuations in vibration information, thus affecting the accuracy of the test results.

[0284] To ensure the reliability of the test results, the control equipment analyzes the superimposed signal after the vibration information stabilizes. Specifically, the control equipment acquires the frequency domain response of the aircraft structural modes at that frequency point based on the excitation signal and the stabilization control signal. By analyzing the frequency domain response, the amplitude and phase information of the aircraft structural modes at that frequency point can be calculated. These amplitude and phase information are then compared with preset amplitude and phase ranges. If the amplitude or phase information is found to be outside the preset range, the frequency and / or amplitude of the excitation signal need to be adjusted. This adjustment aims to optimize the test conditions to more accurately simulate the actual flight environment or meet specific test requirements, thereby ensuring the accuracy and validity of the test results. It should be noted that the preset amplitude and phase ranges are determined based on GVT.

[0285] It should be noted that when the aircraft structure control coupling test is in the open-loop test state, it is also necessary to determine whether the frequency and / or amplitude of the excitation signal needs to be adjusted. Specifically, to ensure that the excitation signal output by the flight controller is consistent with the excitation signal configured by the control equipment, the control equipment needs to detect the excitation signal. The control equipment will determine whether the excitation signal output by the flight controller is the excitation signal configured by the control equipment. When the control equipment confirms that the excitation signal output by the flight controller is consistent with the configured excitation signal, it connects the flight controller to each actuator and begins to acquire vibration information. If it is confirmed that the excitation signal output by the flight controller is inconsistent with the configured excitation signal, then step S602 is executed.

[0286] S602: If it is determined that the frequency and / or amplitude of the excitation signal needs to be adjusted, a parameter adjustment command is sent to the flight controller.

[0287] In this step, based on step S601, if it is determined that the frequency and / or amplitude of the excitation signal needs to be adjusted, the control device sends a parameter adjustment command to the flight controller to optimize the test conditions.

[0288] The parameter adjustment command instructs the flight controller to adjust the frequency and / or amplitude of the excitation signal. This command includes specific adjustment parameters, such as the new frequency value and amplitude change. These parameters are derived from the control equipment's analysis of the current test data and are designed to guide the flight controller in making necessary adjustments to the excitation signal.

[0289] S603: Receives parameter adjustment instructions sent by the control equipment.

[0290] In this step, based on step S602, the flight controller receives a parameter adjustment command sent by the control equipment. This parameter adjustment command instructs the flight controller to adjust the frequency and / or amplitude of the excitation signal.

[0291] S604: Adjust the frequency and / or amplitude of the excitation signal according to the parameter adjustment command.

[0292] In this step, based on step S603, the flight controller adjusts the frequency and / or amplitude of the excitation signal according to the parameter adjustment command.

[0293] Specifically, after receiving parameter adjustment instructions from the control equipment, the flight controller begins to precisely adjust the frequency and / or amplitude of the excitation signal.

[0294] It is important to note that before beginning precise adjustment of the excitation signal's frequency and / or amplitude, the control equipment must preemptively halt excitation signal generation to ensure the safety and accuracy of the subsequent adjustment process. Specifically, halting excitation signal generation prevents unnecessary disturbances to the aircraft during adjustment, avoiding abnormal dynamic responses caused by sudden changes in excitation signal parameters. Simultaneously, halting excitation signal generation provides a stable initial state for the adjustment process, allowing new frequency and amplitude parameters to be accurately set in an interference-free environment. This preemptive halting step enables test personnel to better control the signal adjustment process, reduce potential risks, and improve the overall reliability of the test.

[0295] It is worth noting that the generation of the excitation stop signal can be achieved by the control device sending an excitation stop command to the flight controller. For the specific implementation method, please refer to steps S701 to S703, which will not be elaborated here.

[0296] In one possible implementation, the flight controller first parses the parameter adjustment command. Specifically, the flight controller reads the specific adjustment parameters contained in the parameter adjustment command, such as the new frequency value or amplitude change.

[0297] Next, the flight controller's excitation signal processing module modifies the current excitation signal as necessary to obtain the adjusted excitation signal. If the parameter adjustment command requires frequency adjustment, the flight controller will change the periodicity characteristics of the excitation signal, such as adjusting the sampling rate or recalculating the spectral characteristics of the excitation signal, to ensure that the new frequency can be accurately applied to the aircraft's actuators. For amplitude adjustment, the flight controller achieves this by changing the amplitude coefficient of the excitation signal, thus ensuring that the frequency and amplitude of the excitation signal meet the new test requirements.

[0298] Finally, the adjusted excitation signal will be superimposed on the stability augmentation control signal again to form a comprehensive input signal, which will be applied to the aircraft's actuators.

[0299] The aircraft structure control coupling test method provided in this application can more accurately simulate the actual flight environment by adjusting the frequency and / or amplitude of the excitation signal according to the parameter adjustment command, thereby improving the flexibility and adaptability of the test, ensuring that specific design standards and test objectives are met, and improving the accuracy and reliability of test data. This provides important support for aircraft design optimization and performance verification.

[0300] Figure 7 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 6 ,like Figure 7 As shown, based on any of the above embodiments, the method further includes:

[0301] S701: Sends an excitation abort command to the flight controller. The excitation abort command is used to instruct the flight controller to stop generating an excitation signal.

[0302] In this step, structural control coupling tests of aircraft typically involve applying excitation signals to test the structural response of the aircraft and the reaction of its control system under different conditions. These excitation signals, under certain test conditions, help analyze the dynamic characteristics, stability, and responsiveness of the aircraft. However, as the test progresses, especially when test conditions change, test objectives are adjusted, or the aircraft's state becomes uncertain, it is necessary to promptly stop the generation of excitation signals to ensure the safety and accuracy of the test or to avoid overloading the aircraft.

[0303] During the test, the control equipment needs to monitor the aircraft's status and the test progress in real time. When a change in the aircraft's status is detected or the test process needs adjustment, the control equipment can issue an excitation stop command based on the current test requirements. The issuance of this command depends not only on the operator's instructions but also on the aircraft's real-time status, the control system's feedback, and environmental factors. For example, if the aircraft enters an unstable state or an anomaly occurs during the test, the excitation stop command can be triggered immediately, thereby stopping the application of excitation signals and preventing further interference to the aircraft.

[0304] S702: Receives an excitation stop command sent by the control device.

[0305] In this step, the flight controller receives the excitation stop command sent by the control equipment.

[0306] Specifically, the flight controller first receives the excitation abort command from the control device via a data transmission interface. The flight controller's hardware and software architecture parses and verifies this command to ensure its integrity and validity. Optionally, verification methods include checking the command's format, its timeliness, and whether the command parameters match the current aircraft state. This process ensures that only valid excitation abort commands are processed, preventing misoperation or invalid commands from causing aircraft malfunctions.

[0307] S703: Stop generating excitation signals according to the excitation stop instruction.

[0308] In this step, based on step S702, the flight controller stops generating the excitation signal according to the excitation abort command.

[0309] Specifically, when the flight controller successfully receives the excitation abort command, it immediately begins to execute the operation to stop the generation of excitation signals. Excitation signals are typically generated and sent by the flight controller. If the aircraft is currently executing a predetermined excitation signal, that signal will be immediately interrupted at this point.

[0310] The specific stopping mechanisms include cutting off the excitation source, disconnecting the signal link, and providing real-time feedback. Specifically, cutting off the excitation source means that the source generating the excitation signal is turned off or reset. Disconnecting the signal link means interrupting the transmission channel of the excitation signal, ensuring that the excitation signal is no longer transmitted to the aircraft's actuators. Real-time feedback means that after the excitation signal stops, the flight controller sends feedback information to the control equipment, informing them that the excitation signal has been successfully stopped and that the aircraft's state has returned to normal control mode.

[0311] In one specific implementation, a switch 2 is provided on the output path of the excitation signal. Switch 2 is located before the excitation signal and the stabilization control signal are superimposed. Switch 2 is used to control the input or output of the excitation signal. The control device manages the state of switch 2, determining its on / off state based on the system's operating status and experimental requirements. Under normal operating conditions, switch 2 remains closed, meaning the excitation signal can be transmitted unimpeded to the next stage for superposition with the stabilization control signal. However, under certain specific experimental conditions or emergency situations, such as when system instability is detected, special experimental procedures are required, or to prevent the excitation signal from adversely affecting the system, the control device can quickly disconnect switch 2. By disconnecting switch 2, the system can immediately interrupt the transmission of the excitation signal, thus preventing the excitation signal from affecting system stability at inappropriate times.

[0312] The aircraft structure control coupling test method provided in this application can monitor the aircraft's state changes in real time during the test and promptly stop the generation of excitation signals as needed. This mechanism effectively avoids excessive response or instability of the aircraft that may be caused by continuous excitation signals, ensuring that the aircraft is always in a controllable state. By precisely stopping the excitation signals, not only is the safety of the aircraft improved, but measures can also be taken quickly in the event of abnormal or unforeseen circumstances to prevent the test risks from escalating, thereby ensuring the safety of the aircraft and test personnel.

[0313] The specific implementation process of step S206 will be discussed below. Figure 8 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 7 ,like Figure 8 As shown, based on any of the above embodiments, the method for determining whether the structural control coupling test passes based on vibration information includes:

[0314] S801: Based on vibration information, obtain the excitation response of the aircraft during the test.

[0315] In this step, the control equipment acquires the excitation response of the aircraft during the test based on the vibration information obtained from the inertial navigation and vibration acquisition equipment, laying the foundation for determining whether the structural control coupling test is successful. Excitation response refers to the dynamic reaction exhibited by the aircraft's airframe structure when subjected to an excitation signal.

[0316] Specifically, vibration information includes multi-dimensional data provided by vibration acquisition equipment and inertial navigation, which reflects the structural vibration characteristics of the aircraft under different operating conditions.

[0317] By applying signal processing techniques such as Fourier transform, time-frequency analysis, and modal analysis, control equipment can extract characteristic parameters of the excitation response from vibration information. These parameters include, but are not limited to, natural frequency, damping ratio, and modal shape, which describe the dynamic behavior of the aircraft structure under excitation signals. By analyzing these parameters, test personnel can understand how the aircraft responds to excitation signals during testing and whether this response is within the expected design range.

[0318] S802: Determine whether the structural control coupling test is passed based on the excitation response.

[0319] In this step, based on the excitation response obtained in step S801, the control device determines whether the structural control coupling test has passed.

[0320] Specifically, the control equipment compares the extracted excitation response parameters with pre-set standards. Through this comparison, the control equipment can determine whether the aircraft's structural response meets design expectations. If the excitation response parameters under all test conditions are within the set standard range, it indicates that the aircraft structure can effectively cope with the excitation under the test conditions, and the structural control coupling test is considered passed. If the excitation response parameters under any test condition exceed the standard range, there may be a structural control coupling problem or design flaw, requiring further analysis and improvement.

[0321] For the specific determination process, please refer to steps S901 to S902, which will not be repeated here.

[0322] The aircraft structural control coupling test method provided in this application analyzes vibration information to obtain the excitation response and determines whether the test passes based on the response. This method can accurately assess the dynamic behavior of the aircraft under test conditions, ensuring that its structural performance meets design standards, thereby improving the accuracy and reliability of the test. Simultaneously, this process helps to identify and resolve potential structural problems in a timely manner, supporting the optimization of aircraft design and the improvement of safety.

[0323] The specific implementation process of step S802 will be discussed below. Figure 9 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 8 ,like Figure 9 As shown, based on the above embodiments, the method for determining whether the structural control coupling test passes based on the excitation response includes:

[0324] S901: The structural control coupling test is deemed passed when the excitation response is greater than the preset amplitude-frequency margin index.

[0325] In this step, if the excitation response is greater than the preset amplitude-frequency margin index, the structural control coupling test is considered passed.

[0326] The preset amplitude-frequency margin index is a standard used to measure the dynamic performance of an aircraft airframe structure. It includes the vibration amplitude and frequency response characteristics of the airframe structure at different frequencies; specifically, it involves amplitude margin and frequency margin. Amplitude margin refers to the maximum vibration amplitude the airframe structure can withstand at a specific frequency without performance degradation or instability. Frequency margin relates to the airframe structure's response characteristics at different frequencies, ensuring that the structure maintains stability and performance at these frequencies. By comparing the response with the excitation response, testers can determine whether the structure meets design requirements and make adjustments and optimizations as necessary.

[0327] If the excitation response under all test conditions exceeds the preset amplitude-frequency margin, it means that the aircraft structure exhibits sufficient stability and elasticity under excitation, effectively absorbing and resisting external disturbances. In this case, the control equipment will determine that the structural control coupling test has passed, indicating that the aircraft's structural design meets the expected performance standards and possesses safety and reliability under actual operating conditions.

[0328] S902: The structural control coupling test is deemed unsuccessful when the excitation response is less than or equal to the amplitude-frequency margin index.

[0329] In this step, if the excitation response is less than or equal to the amplitude-frequency margin index, the structural control coupling test is deemed to have failed.

[0330] Specifically, if the excitation response under any test condition is less than or equal to the preset amplitude-frequency margin index, it means that the aircraft structure has failed to exhibit sufficient stability and elasticity under excitation, and cannot effectively absorb and resist external disturbances. In this case, the control equipment will determine that the structural control coupling test has failed. This indicates that the aircraft's structural design may have deficiencies or defects, failing to meet the expected performance standards and posing safety hazards under actual operating conditions.

[0331] In one specific implementation, if the frequency domain response data of each structural mode of the aircraft is less than the preset amplitude-frequency margin index under the open-loop test state, and the time domain response data of each structural mode of the aircraft is less than the preset amplitude-frequency margin index under the closed-loop test state, it means that the open-loop test and closed-loop test of the aircraft meet the requirements. At this time, it can be determined that the aircraft meets the requirements of the aircraft structure-control coupling test and can be put into the market.

[0332] Optionally, this application uses the angular rate signal and acceleration signal of the aircraft as the time-domain response. If the time-range decay of each angular rate signal and each acceleration signal of the aircraft is less than the preset amplitude-frequency margin index, then the time-domain response is considered to meet the requirements.

[0333] It should be noted that the time-domain response is not limited to angular rate and acceleration signals. Other information besides angular rate and acceleration signals can also be used, as long as it is convenient for observation and recording. No restrictions are imposed here.

[0334] It is understandable that the amplitude margin index is not fixed. For different types of aircraft, those skilled in the art can adapt and set the corresponding amplitude margin index, which is not limited here.

[0335] The aircraft structure control coupling test method provided in this application determines whether the structure control coupling test passes based on the excitation response and amplitude-frequency margin index. It not only provides a clear standard to judge whether the structure control coupling test passes or fails, but also provides an important reference for the optimization of structural design. By ensuring that the excitation response exceeds the preset amplitude-frequency margin index, the design of the aircraft can better meet the dynamic performance requirements in actual use.

[0336] In one possible implementation, the amplitude margin specification is 6dB or 9dB.

[0337] Specifically, the amplitude margin is 6dB or 9dB, depending on the stage of the aircraft's lifespan and its required performance.

[0338] In one possible implementation, determining whether the structural control coupling test passes based on the excitation response includes:

[0339] In the open-loop test, if the excitation response under single gain is greater than the preset first amplitude margin index, and the excitation response under double gain is greater than the second amplitude margin index, then the structural control coupling test is deemed to have passed.

[0340] Otherwise, the structural control coupling test is deemed unsuccessful;

[0341] Among them, the first amplitude margin index is greater than the second amplitude margin index.

[0342] Specifically, the excitation response refers to the swept frequency response at all structural modal frequencies.

[0343] In one specific implementation, under open-loop testing conditions, for a newly manufactured aircraft that has not yet undergone its maiden flight, if the excitation response of the aircraft under single-gain conditions is greater than a preset first amplitude margin index, and the excitation response under double-gain conditions is greater than a second amplitude margin index, then the structural control coupling test is determined to have passed; otherwise, the structural control coupling test is determined to have failed. In this case, preferably, the first amplitude margin index is 9dB, and the second amplitude margin index is 0dB.

[0344] In another specific implementation, under open-loop test conditions, for an aircraft with a flight duration within a preset time period, if the excitation response of the aircraft under single-gain conditions is greater than a preset first amplitude margin index, and the excitation response under double-gain conditions is greater than a second amplitude margin index, then the structural control coupling test is determined to have passed; otherwise, the structural control coupling test is determined to have failed. In this case, preferably, the first amplitude margin index is 6dB, and the second amplitude margin index is 0dB.

[0345] During the above experiment, the length of the preset time period can be set according to the actual experimental situation, and there is no limit to the specific duration. For example, the preset time period can be set to 500 hours.

[0346] In another specific implementation, under closed-loop test conditions, if the closed-loop response of the aircraft does not diverge under both single-gain and double-gain conditions, it indicates that the closed-loop test of the aircraft meets the requirements. Optionally, under closed-loop test conditions, the amplitude margin is 6 dB. It is worth noting that the response should conform to the design value of the closed-loop transfer function after adding a notch filter.

[0347] The structural control coupling test method for aircraft provided in this embodiment offers a standard for evaluating the dynamic response capability of an aircraft under excitation conditions by setting a clear and quantifiable amplitude-frequency margin index. Choosing 6dB or 9dB as the amplitude-frequency margin index ensures that the aircraft has sufficient stability and resilience when responding to external disturbances, thereby effectively avoiding structural damage or performance degradation that may result from excessive vibration. This setting not only verifies the robustness of the aircraft design but also provides specific guidance for further design optimization.

[0348] The specific implementation process of step S205 will be discussed below. Figure 10 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 9 ,like Figure 10 As shown, based on any of the above embodiments, the method for obtaining vibration information of the aircraft's airframe structure during the test, acquired by the aircraft's inertial navigation and vibration acquisition equipment, includes:

[0349] S1001: During the test, receive the first vibration information of the aircraft's airframe structure collected by the vibration acquisition equipment.

[0350] In this step, vibration information is the core basis for determining whether the structural control coupling test is passed in the aircraft structural control coupling test. Therefore, in order to reduce the complexity of the test operation and improve the accuracy of data acquisition, it is particularly important to install vibration acquisition equipment on the aircraft airframe structure in advance.

[0351] The vibration acquisition device is an accelerometer, which is used to acquire acceleration, thereby enabling real-time monitoring of the vibration characteristics of the aircraft under different test conditions.

[0352] Specifically, the pre-installed vibration acquisition equipment can automatically record key parameters such as vibration frequency, amplitude, and phase of the aircraft's structure, forming primary vibration information, and then transmit this information wirelessly or via wired means to the control equipment for analysis and processing.

[0353] S1002: Receives the second vibration information of the aircraft's airframe structure during the test, which is collected by the inertial navigation system and uploaded by the flight control system; wherein the vibration information includes the first vibration information and the second vibration information.

[0354] In this step, ensuring the accuracy and reliability of the test results is crucial in the aircraft structure control coupling test. Therefore, the control equipment in this application not only relies on the first vibration information provided by the vibration acquisition equipment, but also receives second vibration information acquired by the inertial navigation system uploaded from the flight control system. The second vibration information provided by the inertial navigation system includes linear acceleration and angular velocity.

[0355] By receiving the second vibration information, the control equipment can obtain the changes in linear acceleration and angular velocity experienced by the aircraft during the test. This multi-dimensional data acquisition method can more accurately reflect the dynamic behavior and structural response of the aircraft, providing a rich information foundation for the analysis of test results.

[0356] The aircraft structure control coupling test method provided in this application forms comprehensive vibration information by acquiring first vibration information from vibration acquisition equipment and second vibration information from inertial navigation. This dual data acquisition strategy not only significantly improves the accuracy of test results but also enhances the robustness of the test. Even under complex test conditions, the control equipment can identify potential structural problems and performance bottlenecks by comprehensively analyzing data from multiple sources.

[0357] Figure 11 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 10 ,like Figure 11 As shown, based on any of the above embodiments, the method further includes:

[0358] S1101: Send a test termination command to the flight controller. The test termination command is used to instruct the flight controller to disconnect from at least one actuator.

[0359] In this step, during the aircraft structure control coupling test, the flight control system is responsible for controlling the operation of each actuator of the aircraft.

[0360] If an abnormal situation occurs during the test and it needs to be terminated, the control equipment sends a test termination command to the flight controller, instructing the flight controller to disconnect from at least one actuator. The purpose of this operation is to terminate the test promptly and prevent further excitation signals from damaging the aircraft structure.

[0361] Abnormal situations include, but are not limited to, any one or any combination of actuator failure, abnormal sensor data, structural vibration exceeding the expected range, or changes in the external environment. This application does not make specific limitations on these situations.

[0362] S1102: Receive the termination test command sent by the control device.

[0363] In this step, based on step S1101, the flight controller receives a termination test command sent by the control device. The termination test command instructs the flight controller to disconnect from at least one actuator.

[0364] S1103: Disconnect the flight controller from at least one actuator according to the test termination command.

[0365] In this step, based on step S1102, the flight controller disconnects from at least one actuator according to the received termination test command.

[0366] The termination test command contains several key parameters, such as the identifier of the actuator to be disconnected, priority information, and the urgency of the disconnection operation.

[0367] In one specific implementation, the flight controller first parses the parameters in the termination test command to determine which actuators need to be disconnected. Then, based on the priority information in the command, the flight controller determines the order and urgency of the disconnection operations. After identifying the actuators that need to be disconnected, the flight controller cuts off the power source or signal channel of these actuators via control signals, ensuring they immediately cease operation. The entire process needs to be completed within a very short time to avoid any potential structural damage or safety risks. Through this precise and rapid response mechanism, the flight controller can effectively protect the structural integrity of the aircraft and provide conditions for subsequent fault analysis and system recovery.

[0368] In one specific implementation, a switch 1 is provided between the flight controller and at least one actuator, and a control device controls the on / off state of switch 1. Specifically, the main function of switch 1 is to control the on / off state of signals. It is managed by the control device to ensure that signals can be transmitted smoothly under normal operating conditions. However, in the event of an emergency, the control device can quickly disconnect switch 1. This function is designed to protect the aircraft's structure in emergency situations and prevent potential damage caused by overload or other abnormal conditions.

[0369] The aircraft structure control coupling test method provided in this application significantly improves the safety and reliability of the test process by timely sending and receiving test termination commands. When an abnormal situation is detected, the control equipment can quickly instruct the flight controller to disconnect from the actuator, thereby immediately stopping potentially dangerous operations and preventing damage to the aircraft structure. This mechanism not only protects the structural integrity of the aircraft but also ensures the safety and stability of the aircraft under various test conditions.

[0370] Figure 12 A flowchart illustrating a flight vehicle structure control coupling test method provided in this application. Figure 10 First, such as Figure 12 As shown, based on any of the above embodiments, the method further includes:

[0371] S1201: After completing the test of the aircraft in any configuration, send a configuration switching command to the flight control system. The configuration switching command is used to instruct the flight control system to switch the aircraft to other configurations for structural control coupling tests.

[0372] In this step, the aircraft structural control coupling test includes tests under different aircraft configurations to ensure the structural integrity and stability of the control system under various configurations. By conducting tests under multiple configurations, the performance of the aircraft under different operating conditions can be verified, and existing structural coupling problems can be identified.

[0373] Therefore, after completing the test of the aircraft in any configuration, the control equipment sends a configuration switching command to the flight controller. This configuration switching command instructs the flight controller to switch the aircraft to another configuration for structural control coupling testing.

[0374] It is worth noting that aircraft are divided into fixed-configuration aircraft and variable-configuration aircraft. Among them, variable-configuration aircraft can be further divided into takeoff and landing configuration, transition configuration, and cruise configuration.

[0375] Specifically, takeoff and landing configuration refers to the configuration in which an aircraft climbs from a stationary ground using vector power and descends using dynamic vectoring. The main actuators of the aircraft are direct force actuators (generally rotors, ducted fans, etc.).

[0376] Transition configurations refer to the transition modes between rotor and fixed-wing modes. These include tilt-rotor transition configurations and dual-rotor transition configurations.

[0377] Cruise configuration refers to the configuration that, after completing the entire transition phase, allows the aircraft to cruise using lift provided by its wings, much like a fixed-wing aircraft. The primary control surfaces of the aircraft are ailerons, elevators, rudders, or other equivalent mechanisms (elevators, rudders, etc.).

[0378] It should be noted that the configuration breakdown of variable-configuration aircraft also includes the configuration of aircraft equipment status. Because the configuration of aircraft equipment status encompasses a wide range, such as the number of battery packs, and because different battery packs have a significant impact on the aircraft's weight, it will be managed separately in the experimental design.

[0379] Furthermore, in the structural control coupling test of the aircraft, a support device is needed to support the aircraft in different configurations.

[0380] In one possible implementation, the support device includes a flexible suspension system or a support frame. The flexible suspension system is used to suspend the aircraft for open-loop and closed-loop tests of takeoff and landing configurations, transition configurations, and cruise configurations. The support frame is used to support the aircraft during open-loop and closed-loop tests of takeoff and landing configurations.

[0381] S1202: Receive configuration switching instructions sent by the control device.

[0382] In this step, based on step S1201, the flight controller receives the configuration switching command sent by the control device.

[0383] Among them, the configuration switching command is used to instruct the flight control system to switch the aircraft to other configurations for structural control coupling tests.

[0384] S1203: According to the configuration switching command, control the aircraft to switch to other configurations to conduct structural control coupling tests.

[0385] In this step, based on step S1202, the flight controller controls the aircraft to switch to other configurations according to the configuration switching command to conduct structural control coupling tests.

[0386] In one specific implementation, upon receiving a configuration switch command, the flight controller initiates a preset control program to achieve the aircraft's configuration transition. First, the flight controller comprehensively assesses the current flight status of the aircraft to ensure all aircraft parameters are within safe ranges, facilitating a smooth configuration switch. Next, the flight controller gradually adjusts the aircraft's control surfaces, such as the elevator and rudder, to alter the aircraft's aerodynamic characteristics and adapt it to the new configuration requirements.

[0387] During the switchover, the flight control system continuously monitors key parameters such as speed, altitude, attitude angles, and acceleration to ensure the aircraft remains stable when transitioning to the new configuration. Any anomalies are detected in real time and trigger corresponding emergency procedures to ensure the aircraft's safety.

[0388] Once the configuration switch is complete, the flight control system places the aircraft in the new configuration and begins structural control coupling tests. These tests aim to evaluate the aircraft's structural response and control system performance under the new configuration, ensuring its reliability and safety under various operating conditions.

[0389] The structural control coupling test method for aircraft provided in this application, after completing tests of the aircraft in any configuration, sends a configuration switching command to the flight control system and receives and executes the command, thereby enabling structural control coupling tests of the aircraft in different configurations. This method can systematically evaluate the structural integrity and control system stability of the aircraft under multiple configurations. By conducting tests under multiple configurations, potential structural coupling problems can be identified and analyzed, improving the reliability and safety of the aircraft design.

[0390] In one possible implementation, during the closed-loop test, the notch filter is controlled to adjust the stabilization control signal based on the amplitude and phase information of the excitation signal and the stabilization control signal.

[0391] In this step, if the current test is in a closed loop, the notch filter needs to be adjusted to control the stabilization control signal based on the amplitude and phase information of the excitation signal and the stabilization control signal, in order to identify and suppress unwanted vibration frequencies, thereby optimizing the stability and response performance of the aircraft.

[0392] In one specific implementation, during the initial stage of the closed-loop test, under the action of the excitation signal, a pre-designed notch filter is used to filter the vibration information collected by the inertial navigation and vibration acquisition equipment to obtain the initial stabilization control signal. Then, the stabilization control signal and the excitation control signal are superimposed, and a preset allocation matrix is ​​used to distribute the superimposed signal to each actuator, obtaining the control quantity on each actuator. This allows the actuators to control the aircraft, while the inertial navigation and vibration acquisition equipment collects the aircraft's vibration information.

[0393] After the vibration information is transmitted back to the control equipment, the control equipment instructs the flight controller to filter the vibration information using a notch filter, obtaining filtered data. Based on this filtered data, the stabilization control signal is adjusted. This process first requires spectral analysis of the excitation and stabilization control signals to obtain their amplitude and phase information. Amplitude information helps determine the signal strength, while phase information is used to identify the signal's timing characteristics. By analyzing this information, frequency components that may cause instability under the current test conditions can be identified.

[0394] Once these frequencies are identified, the notch filter can be adjusted to precisely suppress these specific frequencies. The filter's center frequency, bandwidth, and notch depth can be adjusted based on amplitude and phase information to ensure effective suppression of unwanted vibration frequencies. The adjusted notch filter then processes the stabilization control signal, filtering out unstable frequencies and thus optimizing signal stability.

[0395] The adjusted stabilization control signal and excitation control signal are superimposed, and the above process is repeated. Through this dynamic feedback loop, the notch filter is continuously adjusted as the experiment progresses to adapt to different test conditions and response characteristics.

[0396] It should be noted that vibration information can be obtained through sensors on the inertial navigation system, sensors on the vibration acquisition device, or a combination of inertial navigation and vibration acquisition devices; no limitation is made here.

[0397] The aircraft structure-control coupling test method provided in this embodiment analyzes the amplitude and phase information of the excitation signal and the stabilization control signal to accurately identify and adjust the notch filter, thereby suppressing specific vibration frequencies that may occur in different aircraft configurations. This dynamic adjustment process significantly improves the stability and control accuracy of the aircraft. This method not only ensures the reliable performance of the aircraft under various operating conditions but also provides solid data support for optimized design and improved overall safety.

[0398] It is worth noting that for rotorcraft or compound wing aircraft, when conducting structural control coupling tests, a high-energy rotating propeller must be driven to effectively excite the control channels, thereby stimulating the modal characteristics of the aircraft structure. To ensure that the test conditions are consistent with the control excitation states during ground takeoff and in-flight, the propeller needs to operate at the same speed as in actual flight.

[0399] However, applying such high rotational speeds while the aircraft is supported by landing gear on the ground or suspended in the air can pose significant safety risks and may even be impossible in some situations. Therefore, in the structural control coupling tests of rotor-configured aircraft, a suitable propeller must be carefully selected as the excitation source to ensure the safety and effectiveness of the test.

[0400] In one possible implementation, the propeller blades of a rotorcraft or compound wing aircraft can be used as the excitation source, although this requires reducing the collective pitch for aircraft with larger propellers. This adjustment is achieved through aerodynamic calculations to ensure that the thrust generated by the blades is within the set thrust requirements, so that appropriate thrust or pull can still be generated under different test conditions, thereby effectively stimulating the structural characteristics of the aircraft.

[0401] However, if reducing the collective pitch to the minimum still fails to meet the set tension requirements, then replacing the propeller needs to be considered. The reason for replacing the propeller is not only to meet the tension requirements, but also because interference may occur between the suspended rubber cable and the propeller, posing a certain danger. Replacing the propeller with a suitable one can effectively avoid this potential interference risk, ensuring the safety and smooth conduct of the experiment.

[0402] In one possible implementation, if the propeller needs to be replaced, the propeller of the rotorcraft or compound wing aircraft can be selected according to the following formula (1) or according to the following formulas (1) and (2):

[0403]

[0404] Where ρ is the atmospheric density, N is the number of propeller blades, π is pi, and C L R is the propeller thrust coefficient, R is the propeller radius, L is the propeller occupancy information, and N is the propeller's thrust coefficient. RPM d is the nominal rotational speed of the propeller. RPM f is the maximum value of the propeller differential speed. min The lowest sweep frequency is I, where I is the inertia of the aircraft under test, and p is the lowest sweep frequency. max The maximum angular velocity driving the aircraft is M, the weight of the aircraft under test is M, and α is a preset proportional threshold. Preferably, α is 0.2.

[0405] By selecting a suitable small propeller as the excitation source to replace the large propeller, and by giving the selection principle of the small propeller as the excitation source, the experimental method is improved and the experimental safety is greatly enhanced.

[0406] For rotorcraft or compound aircraft, open-loop tests are conducted using the selected propeller as the excitation source. For fixed-wing aircraft, open-loop tests can be conducted directly. Based on the open-loop structure control coupling test results of each channel and the preset design parameters of the notch filter, a second-order notch filter is designed according to formula (3), and the parameters of the notch filter are adjusted according to the open-loop test data. Formula (3) is shown below:

[0407]

[0408] Where s is the Laplace transform factor, a and b are the parameters to be adjusted in the notch filter, and w n The natural frequency is used. Specifically, if the test data at the current frequency does not meet the first standard, the parameters of the notch filter are adjusted until the test data at the current frequency meets the requirement that the frequency domain response data of each structural mode of the aircraft is less than the preset amplitude margin index. Then, the test data at the next frequency is obtained, and the above-mentioned notch filter parameters are adjusted again.

[0409] Below, we will take the eVTOL aircraft as an example to explain in detail the problems existing in the current technology.

[0410] Specifically, Figure 13 This application provides a modal diagram of a conventional fixed-wing aircraft wing with a single bend. Figure 14This application provides a modal diagram of a single-bend structure for an eVTOL wing. For compound wing aircraft, especially emerging types like eVTOLs, most eVTOL aircraft employ distributed propulsion systems. Unlike fixed-wing aircraft, distributed propulsion systems typically use arm structures added to the wing to distribute the power units at both ends. This layout significantly increases the mass distribution on the outer section of the wing. However, due to level flight performance requirements, the airfoil stiffness cannot increase proportionally to the mass distribution. Therefore, this mass-distributed arrangement leads to increased inertia, resulting in a decrease in the structural modal frequency.

[0411] Meanwhile, due to the dispersed layout of large mass points in the distributed propulsion system of eVTOL aircraft, the number of inherent modes is significantly increased compared to fixed-wing aircraft, and the mode distribution is very dense. The distributed propulsion excitation sources are located at 1 / 4, 1 / 5, and 1 / 6, with the antinodes and nodes of the main mode wave, making it easy to excite higher-order modes. Furthermore, the motor speed is significantly lower than that of traditional engines, which makes the frequency of the propulsion control excitation source closer to the frequency of the structural modes, thus greatly increasing the risk of coupling between the overall structural modes and servo control. Due to the complex structural design and distributed propulsion layout of eVTOL aircraft, the number of structural modes is far greater than that of fixed-wing or helicopter aircraft. Besides typical symmetrical and antisymmetrical structural modes, such as... Figures 15 to 20 As shown, a number of new modes induced by the arm were also generated, such as arm-symmetric, anti-symmetric pitch and swing structural modes.

[0412] in, Figure 15 This application provides a modal diagram of a symmetrical pitch structure for an eVTOL external arm. Figure 16 This application provides a modal diagram of an anti-symmetric pitch structure of an eVTOL external arm. Figure 17 This application provides a modal diagram of a symmetrical pitch structure of an eVTOL internal arm. Figure 18 This application provides a modal diagram of an anti-symmetric pitch structure of an eVTOL internal arm. Figure 19 This application provides a modal diagram of a symmetrical swing structure of an eVTOL internal arm. Figure 20 Modal diagram of an anti-symmetric swing structure of an eVTOL internal arm provided in this application. Based on Figures 15 to 20As shown, the influence mechanisms of these new structural modes differ significantly from those of traditional modes. This is especially true for the massive lifting rotor connected to the arm, where the additional velocity and deformation induced by the structural modes are further coupled with changes in the control rotor vector direction and efficiency. This makes traditional structural modal simulation analysis methods and tools inadequate for current experimental needs, and there is also a lack of corresponding simulation and analysis tools and methods for the coupling theory of servo control and structural modes. Addressing these practical difficulties, the inventors have proposed an aircraft structural control coupling test method, apparatus, and equipment that effectively solves these problems, enabling the aircraft to meet the relevant requirements of civil aviation airworthiness regulations (see Articles 23.251, 27.241, and 27.251 of the Civil Aviation Airworthiness Regulations).

[0413] It is worth noting that the control equipment is equipped with a host computer, which can be used to conduct aircraft structure control coupling tests. Figure 21 This is a schematic diagram of the host computer interface for an aircraft structure control coupling test provided in this application. Figure 21 As shown, the functions of this host computer include, but are not limited to:

[0414] At the start of the aircraft structure control coupling test, the excitation type is selected. The excitation type includes single-point excitation or frequency sweep excitation. For single-point excitation, the frequency and number of sweeps of the single-point excitation signal can be set. For frequency sweep excitation, the initial frequency, ending frequency, sweep step size, and number of sweeps of the excitation signal can be set. Simultaneously, data such as the amplitude and normal virtual force of the excitation signal can also be set to obtain control quantities acting on different actuators through an allocation matrix.

[0415] By setting up channels for adding excitation signals, such as pitch, yaw, or roll channels, simulations of different flight states can be achieved.

[0416] By setting attitude protection values, data of the aircraft under critical conditions can be simulated, and the control equipment will automatically shut down the test if the vibration exceeds the limit.

[0417] By setting whether the excitation signal is on or off, the connection and disconnection of the excitation signal can be controlled.

[0418] Optionally, you can also set control closed-loop channels, such as pitch channel, yaw channel or roll channel, and you can choose at least one of them or none of them.

[0419] After the settings are complete, turn on the start switch. The control device will automatically send the set data to the flight controller's onboard software via network communication, and the flight controller will then control the aircraft to simulate flight.

[0420] After the onboard software receives vibration information from each controlled object on the aircraft, it transmits the vibration information back to the control equipment. The main control equipment determines whether the vibration information meets the test requirements. If it does, the corresponding button changes color; otherwise, it remains unchanged. Furthermore, for ease of observation, the host computer displays the real-time operating status of the onboard software, including but not limited to "running" and "stopped." In addition, during the test, the control equipment records the test data in real time and saves it to a database for later retrieval.

[0421] Figure 22 This is a schematic diagram of an aircraft structure control coupling test system provided in this application. Figure 22 As shown, the aircraft structure control coupling test system includes a flight controller 102, a main control PC 220, an inertial navigation system 2201, an aircraft structure 2202, an actuator 2203, and a vibration acquisition device 2204.

[0422] The aircraft structure 2202 can be a fixed-wing aircraft, a rotary-wing aircraft, or a compound-wing aircraft; this application does not make any specific limitations on this.

[0423] The aircraft structure 2202 is the controlled object, and actuators, namely inertial navigation 2201, actuator 2203 and vibration acquisition equipment 2204, are installed on its airframe structure.

[0424] It is worth noting that if the aircraft type is a fixed-wing aircraft, the actuators of this type of aircraft are mainly servos installed in the elevator, rudder and aileron positions, or servos installed in the elevator, rudder and aileron positions.

[0425] If the aircraft type is a rotorcraft, then the actuators for this type of flight are mainly electric motors mounted on the arms to drive the propellers.

[0426] If the aircraft type is a compound wing aircraft, then the actuators of this type of flight are mainly servo motors installed in the elevator, rudder and aileron positions, and motors installed on the arms to drive the propellers; or the actuators of this type of aircraft are mainly servo motors installed in the elevator, rudder and aileron positions, and motors installed on the arms to drive the propellers.

[0427] Specifically, if the fixed-wing aircraft is a V-tail aircraft, then the actuators for this type of aircraft are mainly the servos mounted on the elevator and ailerons of the V-tail. If the compound-wing aircraft involves the elevator design of a V-tail aircraft, then the actuators for this type of aircraft are mainly the servos mounted on the elevator and ailerons, as well as the motors mounted on the arms to drive the propellers.

[0428] During aircraft operation, the pilot / flight controller uses loop-controlled actuators such as control surfaces and motors (2203) to form a pilot / flight controller closed-loop control system. The control frequencies generated by this system may resonate with the modal frequencies of the airframe structure. Once resonance occurs, it will inevitably lead to rapid damage or even disintegration of the airframe structure.

[0429] During the test, the aircraft structure 2202 requires a support device for load bearing, and the frequency of the support device must be lower than the frequency threshold of the first-order structural modal frequency of the aircraft. Preferably, the frequency threshold is one-third of the first-order structural modal frequency of the aircraft.

[0430] The actuator 2203 communicates directly with the flight controller 102 via a serial port / network port and receives control command signals sent by the flight controller 102.

[0431] The inertial navigation system 2201 also communicates directly with the flight control system 102 via a serial port, feeding back the measured aircraft status and vibration information to the flight control system 102.

[0432] Vibration acquisition equipment 2204 is installed at different locations on the aircraft fuselage. Optionally, the specific location of the vibration acquisition equipment 2204 on the aircraft fuselage can be determined according to the requirements of the ground vibration test. It is worth noting that vibration acquisition equipment 2204 must be installed at the mounting locations of inertial navigation system 2201 and motors to allow test personnel to collect and analyze vibration information of the aircraft structure.

[0433] The flight controller 102 is connected to the actuator 2203, the inertial navigation system 2201, and the main control PC 220.

[0434] The Flight Control 102 consists of two parts: computer hardware and software. The software includes flight control software and onboard software. The flight control software provides stability augmentation control laws, enabling the aircraft to achieve stability augmentation control in closed-loop mode. The allocation matrix, part of the flight control software, is used to allocate stability augmentation control laws and frequency sweep signals, converting virtual force and torque control quantities on the lateral, longitudinal, and directional control channels into control quantities for various actuators, specifically including fixed-wing control surface commands, rotorcraft motor throttle commands, and helicopter pitch control servo commands. The onboard software provides frequency sweep signal generation functionality, enabling the generation of experimental frequency sweep signals.

[0435] The main control PC220 is connected to the flight controller 102 via a network port. The main control PC220 has host computer software installed, which allows configuration of the signal waveform information generated by the airborne software. Specifically, without adding any test equipment, the host computer software can be used to adjust the generation, termination, frequency, and amplitude of the sweep signals in the airborne software of the flight controller 102, and simultaneously collect and monitor aircraft vibration information acquired by the inertial navigation system 2201.

[0436] In addition, the main control PC220 is also connected to the vibration acquisition device 2204, which can directly monitor and analyze the vibration information collected by the vibration acquisition device 2204. At the same time, the main control PC220 also has attitude limit measures to ensure that it can automatically shut down after the vibration exceeds the limit.

[0437] It should be noted that before the aircraft structure control coupling test begins, communication tests between the flight controller 102, inertial navigation system 2201, motors, electronic control system, vibration acquisition equipment 2204, and main control PC 220 need to be completed.

[0438] It should be noted that a test task statement is required when conducting aircraft structural control coupling tests. This test task statement is provided by the party commissioning the test. The content of the test task statement includes, but is not limited to, the test name, test purpose, test items, test aircraft status, test conditions, aircraft support requirements, test requirements, layout requirements, and necessary instructions.

[0439] Meanwhile, the test outline shall be prepared by the test undertaking party in accordance with the test mission statement and airworthiness requirements, and shall be approved by the relevant department. The test outline shall generally include the source of the mission, the name of the test, the basis of the test, the purpose of the test, the status of the test aircraft, the test status, the test items, the aircraft support, the test methods and basic principles, the test equipment, the layout of the measurement points, the test procedures, the test safety protection and quality assurance measures, and the principles and contingency plans for handling major problems at the test site.

[0440] During testing, the structure and all systems should be fully loaded, meet all loading and counterweight requirements in the test mandate, pass inspection, and undergo airworthiness manufacturing conformity inspection.

[0441] Dedicated equipment should be used for aircraft structural control coupling tests. The test equipment includes aircraft support equipment, measuring equipment, vibration equipment, etc. The equipment and instruments used should meet the needs of the test, and all equipment and instruments should be in good working order. All measuring instruments must be qualified by metrology verification / calibration and within their validity period, and corresponding reports or verification / calibration certificates should be provided.

[0442] The test process is executed according to the pre-set test control procedure, and relevant test documents are prepared as needed. Test conditions include completing the test outline, determining the test plan, confirming the aircraft configuration, confirming the test environment, installing test equipment according to the test outline, completing communication tests between flight control, inertial navigation, motors and electronic controls, vibration acquisition equipment and control equipment, and collecting vibration signals from the inertial navigation system and vibration acquisition equipment installed at the same location. The test procedure should include the specific test environment, detailed test cases, and specific experimental execution steps. A typical aircraft structure control coupling test procedure includes completing test preparation, ensuring proper aircraft support, starting the frequency sweep excitation test according to the test plan, ensuring the input excitation signal is the desired value, processing the input and output signals to obtain the frequency response function, and ending the test if the result meets the requirements (reasonable amplitude margin); if not, adjusting the notch filter parameters and retesting the affected frequency points until the requirements are met.

[0443] It should be noted that during the structural control coupling test of an aircraft, problems such as poor consistency of test results, incompatibility between command output frequency response characteristics and sensor response, or large differences between data at a certain frequency point and theoretical simulation values ​​are often encountered. Typical causes are: abnormal aircraft status, test equipment problems, or test method problems.

[0444] Among these, abnormal aircraft condition leading to test abnormalities refers to situations where abnormal aircraft condition causes significant changes in test results. If the aircraft's airframe structure, support status, or loadout status does not conform to the test requirements, the vibration modes of the airframe will change significantly during the test. If the fly-by-wire flight control system's operating status does not conform to the test requirements, the sensor frequency response curve will not be affected, but the command output frequency response curve will change significantly. If the control stick circuit is not disconnected during the test, disturbances generated by the control stick will significantly interfere with the test results.

[0445] Test equipment issues leading to abnormal test results refer to situations where abnormalities exist in the cable connections or equipment status of the test equipment, such as unreliable grounding causing high-frequency interference, which can affect the test results. Servo motors exhibit saturated nonlinearity, meaning that their amplitude attenuates significantly and their phase lags noticeably at high frequencies, drastically reducing the servo motor's ability to drive the control surfaces and excite the fuselage vibration. Therefore, the actual vibration value of the fuselage structure will be significantly attenuated at high frequencies, making it more sensitive to high-frequency interference.

[0446] It is worth noting that this application provides normative references. Specifically, the contents of these documents, through normative references within the text, constitute essential provisions of this application. For dated references, only the edition cited applies; for undated references, the latest edition of the referenced document (including any amendments) applies. Normative references include RIN 2120-AL72 Integration ed-Lift Pilot Certification and Operations Miscellaneous Amendments Related to Rotorcraft Airplane Final Rule, GJB 67.7A-2018 Military Aircraft Structural Strength Specification Part 7: Aeroelasticity, and HB8466-2014 General Requirements for Ground Vibration Testing of Civil Aircraft.

[0447] Figure 23 This application provides a schematic diagram of a flight vehicle structural control coupling test system. Figure 23 As shown, the schematic diagram of the aircraft structure control coupling test involves switch 1, switch 2, switch 3, measurement 1, measurement 2, measurement 3, allocation matrix, actuator, controlled object, sensor and stabilization control law.

[0448] Switch 1 is located between the flight controller and at least one actuator, specifically at the position where the sweep frequency control signal (excitation signal) and the stabilization control signal are superimposed. It is used to control the on / off state of all signals. Notably, in an emergency, disconnecting switch 1 can prevent structural damage to the aircraft.

[0449] Switch 2 is located on the output path of the sweep frequency control signal (excitation signal), that is, before the sweep frequency control signal (excitation signal) and the stabilization control signal are superimposed. It is used to control the access or cut-out of the sweep frequency control signal (excitation signal).

[0450] Switch 3 is set on the output path of the stabilization control signal, that is, the position after the sweep frequency control signal (excitation signal) and the stabilization control signal are superimposed, and is used to control the state of the test to be either open-loop or closed-loop.

[0451] Measurement 1 is located after the sweep frequency control signal (excitation signal) and the output of the stabilization control law, and is used to measure and record the control quantity before entering the allocation matrix.

[0452] Measurement 2 is located after the output of the stabilization control law and is used to measure and record the control quantity of the stabilization control law.

[0453] Measurement 3 is located after the sensor and is used to observe and record the sensor measurements.

[0454] The host computer, PC2200, allows for the setting of switch 3 to open and close, thereby enabling the switching on and off of the stability-enhancing control law in the flight control software, achieving open-loop and closed-loop control. Furthermore, the onboard software configured on PC2200 can generate information such as the frequency and amplitude of the excitation signal. PC2200 can also detect whether the excitation signal output by the flight controller meets the requirements from measurement point 1. After confirming that the signal is correct, the connection switch between the flight controller and actuator 2203 is activated.

[0455] By monitoring and comparing the frequency sweep excitation signal acquired at measurement point 1 and the control signal data acquired at measurement point 2 using the main control PC2200, the amplitude of the excitation signal can be adjusted as needed. Once the vibration data waveform stabilizes, the amplitude and phase information of that frequency point are plotted based on the data from measurement points 1 and 2. Then, the notch filter parameters corresponding to that frequency point in the flight control software are adjusted based on this information. Simultaneously, the control and excitation signals from other auxiliary channels are observed, and their amplitude, phase, and frequency response information is plotted.

[0456] It is important to note that when changing the frequency, the excitation signal input switch 1 should be disconnected first. This operation ensures that the new excitation signal is fully modulated and confirmed before injection, and also facilitates data recording for offline analysis.

[0457] Figure 24 This application provides a schematic diagram of a closed-loop test. Figure 24 As shown, the aircraft structure-control coupling test system involves control law G1(s), actuator G2(s), control surface motion G3(s), aircraft elastic structure G4(s), and overload / angular rate sensor G5(s). It is worth noting that... Figure 24 K in the figure is the stability reserve value measured by the open-loop frequency domain test method. Different amplitude pulse excitations are applied to the actuator command input, and the system response (output of inertial sensor, command output of flight control computer and deflection of control surface) is recorded. The aircraft system will not produce constant amplitude oscillations under different excitations; otherwise, it indicates that there is a problem with the open-loop frequency domain test results.

[0458] Figure 25 This application provides a schematic diagram of an open-loop test. Figure 25 As shown, the aircraft structure-control coupling test system involves control law G1(s), actuator G2(s), control surface motion G3(s), aircraft elastic structure G4(s), and overload / angular rate sensor G5(s). Specifically, a swept-frequency signal X is injected into the actuator's command input port to excite it. The flight control computer's command output is the system's open-loop output Y. The frequency domain characteristics of Y / X (represented by a Bode plot) are the open-loop frequency domain characteristics of the structural coupling, from which the stable reserve value can be obtained.

[0459] Figure 26A schematic diagram of the framework of an aircraft structure control coupling test system provided in this application Figure 1 .like Figure 26 As shown, the framework of the aircraft structure control coupling test system includes control equipment 101, vibration acquisition equipment 2204, and aircraft 104. Although Figure 26 Only one control device 101, one vibration acquisition device 2204 and one aircraft 104 are shown in the diagram, but it should be understood that there may be two or more control devices 101, vibration acquisition devices 2204 and aircraft 104.

[0460] Among them, the control device 101 is mainly used to configure the frequency range corresponding to the test in the flight control 102 and to determine whether the structural control coupling test is passed based on the vibration information.

[0461] Optionally, in one specific embodiment, the control device 101 is equipped with a user software interface operated via a touchscreen. When the tester sets the occurrence time, the control device 101 automatically triggers the excitation signal at the predetermined time. When the set termination time is reached, the control device 101 automatically stops the output of the excitation signal. This precise time control allows the tester to flexibly manage the duration of the excitation signal application during the test, in order to collect the required data at specific test stages. Simultaneously, the software interface of the control device 101 also allows the tester to set the frequency and amplitude of the excitation signal. For frequency setting, the tester can choose to input a specific single-point frequency or define a frequency range, including the start frequency, end frequency, number of sweeps, and sweep step size. The control device 101 transmits these setting parameters to the flight controller 102, which generates the corresponding excitation signal according to the settings. If a frequency range is set, the flight controller 102 automatically scans and generates excitation signals of different frequencies within that range to identify the natural frequencies and resonance phenomena of the aircraft structure. Similarly, for amplitude setting, the tester can choose to input a fixed amplitude or an amplitude range. The control device 101 transmits these settings to the flight controller 102, which then generates an excitation signal with the appropriate intensity.

[0462] It is worth noting that the control device 101 can not only configure any one of the parameters of excitation signal generation, termination, frequency and amplitude individually, but also can arbitrarily combine these parameters. This application does not make any specific limitations on this.

[0463] The vibration acquisition device 2204 is mainly used to collect vibration information from the airframe structure of the aircraft 104 during execution. It is worth noting that the vibration acquisition device 2204 can be a single sensor device, such as an accelerometer. The vibration acquisition device 2204 can also be a combination of multiple sensors to achieve more comprehensive vibration data acquisition. This application does not specifically limit the specific form and type of the vibration acquisition device 2204.

[0464] Aircraft 104 is the object for structural control coupling test. For specific details, please refer to step S201. It will not be described again here.

[0465] Specifically, the aircraft 104 is equipped with a flight control system 102, an inertial navigation system (INS), and actuators. The control device 101 is connected to the flight control system 102 of the aircraft 104 and is used to configure the frequency range required for the test. The flight control system 102 generates excitation signals according to this frequency range and outputs them to at least one actuator. The inertial navigation system of the aircraft 104 and the vibration acquisition device 2204 mounted on its airframe structure are responsible for collecting vibration information during the execution process. The control device 101 determines whether the structural control coupling test passes based on this vibration information.

[0466] Figure 27 A schematic diagram of the framework of an aircraft structure control coupling test system provided in this application Figure 2 .like Figure 27 As shown, the framework of the aircraft structure control coupling test system includes a control device 101 and a vibration acquisition device 2204 mounted on the airframe structure of the aircraft 104. The aircraft 104 is equipped with a flight controller 102, an inertial navigation system 2201, and an actuator 2203.

[0467] Specifically, in the aircraft structure control coupling test system, the control device 101 is mainly used to configure the frequency range corresponding to the test in the flight control 102 and to determine whether the structure control coupling test has passed based on the vibration information. The control device 101 includes, but is not limited to, terminal equipment, servers, etc., which are not specifically limited in this application.

[0468] Vibration acquisition device 2204 is installed on the airframe structure of aircraft 104 to collect vibration information on the airframe structure of aircraft 104 during the test.

[0469] Aircraft 104 is the object for structural control coupling test. For specific details, please refer to step S201. It will not be described again here.

[0470] The flight controller 102 is mainly used to receive and process frequency range setting information from the control device 101 and generate corresponding excitation signals to drive the actuator 2203.

[0471] The inertial navigation system 2201 is used to collect vibration information on the airframe structure of the aircraft 104 during the test, and together with the vibration acquisition device 2204, it provides feedback data.

[0472] The actuator 2203 performs corresponding actions based on the excitation signal generated by the flight controller 102 to influence the structural dynamics of the aircraft 104, thereby achieving the objective of the structural control coupling test. It is worth noting that different types of aircraft have different actuators 2203, including but not limited to servos located at the elevator, rudder, or elevator / rudder positions, servos located at the aileron positions, and motors driving the propellers, etc. This application does not specifically limit these.

[0473] The following section will explain the connection methods between the various devices in the aircraft structure control coupling test system.

[0474] The control device 101 is connected to the flight controller 102 via a wired or wireless communication interface to transmit frequency range setting information and determine whether the structural control coupling test is passed based on vibration information.

[0475] The flight controller 102 is connected to the actuator 2203 via a signal line or a dedicated control bus to send excitation signals and control its actions.

[0476] The vibration acquisition device 2204 is connected to the actuator 2203 via a data bus or wireless connection to acquire vibration information of the aircraft 104's airframe structure in real time. Simultaneously, the vibration acquisition device 2204 is connected to the control device 101 via a data bus or wireless connection to transmit the acquired vibration information in real time.

[0477] The inertial navigation system 2201 is installed on the airframe structure of the aircraft 104 and is connected to the flight controller 102 via signal lines or a dedicated control bus to provide real-time vibration information.

[0478] The above connection method can ensure that data and commands can be transmitted efficiently and accurately between various devices in the aircraft structural control coupling test system, thereby supporting the smooth progress of the structural control coupling test.

[0479] The execution logic of the aircraft structure control coupling test system will be explained in detail below.

[0480] First, the control device 101 sets the required frequency range for the test in its configuration interface and transmits this information to the flight controller 102 via a wired or wireless communication interface. After receiving the frequency range setting information, the flight controller 102 generates the corresponding excitation signal and sends it to the actuator 2203 via a signal line or control bus.

[0481] In one possible implementation, the generation of the excitation signal first involves frequency scanning, i.e., a systematic scan within a preset frequency range to ensure coverage of all critical frequencies affecting the aircraft structure. Next, an appropriate signal type is selected based on the specific requirements of the experiment, such as a sine wave for simple periodic vibration testing. Then, the amplitude of the excitation signal is set according to the experimental requirements. The amplitude must be large enough to elicit a noticeable structural response, but not so large as to damage the structure. Once the parameters of the excitation signal are determined, a signal synthesizer or digital signal processor is used to generate the required signal, which is then transmitted to the actuator 2203 via the signal lines or control bus of the flight controller 102. Finally, upon receiving the excitation signal, the actuator 2203 performs corresponding physical actions to influence the structural dynamics of the aircraft 104.

[0482] The actuator 2203 performs specific actions based on the received excitation signals to influence the structural dynamics of the aircraft 104.

[0483] Meanwhile, vibration acquisition device 2204 and inertial navigation system 2201 are respectively installed on the airframe structure of aircraft 104 to collect vibration information in real time. Vibration acquisition device 2204 transmits vibration information to control device 101 via data bus or wireless connection, while inertial navigation system 2201 transmits vibration information to flight controller 102 via signal line or control bus.

[0484] Furthermore, the flight controller 102 transmits vibration information to the control device 101 via a wired or wireless communication interface. The control device 101 determines whether the structural control coupling test is successful based on the vibration information sent by the vibration acquisition device 2204 and the vibration information sent by the inertial navigation system 2201 through the flight controller 102.

[0485] Figure 28 This is a schematic diagram of an embodiment of an aircraft structure control coupling test device provided in this application, as shown below. Figure 28 As shown, the aircraft structure control coupling test device 2800 includes:

[0486] The transmitting module 2801 is used to send an excitation signal generation command to the flight control of the aircraft. The excitation signal generation command is used to instruct the flight control to generate an excitation signal to drive at least one actuator of the aircraft.

[0487] The acquisition module 2802 is used to acquire the vibration information of the aircraft's airframe structure during the test, which is collected by the aircraft's inertial navigation and vibration acquisition equipment;

[0488] The processing module 2803 is used to determine whether the structural control coupling test has passed based on the vibration information.

[0489] In one possible implementation, the sending module 2801 is further configured to:

[0490] Send excitation configuration information to the flight controller. The excitation configuration information carries the frequency range used to generate the excitation signal.

[0491] In one possible implementation, the sending module 2801 is further configured to:

[0492] Send test status control commands to the flight controller. These commands are used to instruct whether to conduct an open-loop test or a closed-loop test.

[0493] In one possible implementation, the processing module 2803 is further configured to:

[0494] After the flight controller generates a stabilization control signal and superimposes it onto the excitation signal, it determines whether to adjust the frequency and / or amplitude of the excitation signal based on the excitation signal and the stabilization control signal.

[0495] In one possible implementation, the sending module 2801 is further configured to:

[0496] If it is determined that the frequency and / or amplitude of the excitation signal needs to be adjusted, a parameter adjustment command is sent to the flight controller, which instructs the flight controller to adjust the frequency and / or amplitude of the excitation signal.

[0497] In one possible implementation, the sending module 2801 is further configured to:

[0498] Send an excitation abort command to the flight controller. The excitation abort command is used to instruct the flight controller to stop generating excitation signals.

[0499] In one possible implementation, the acquisition module 2802 is also used for:

[0500] Based on the vibration information, the excitation response of the aircraft during the test was obtained.

[0501] In one possible implementation, the processing module 2803 is further configured to:

[0502] Based on the excitation response, determine whether the structural control coupling test is successful.

[0503] In one possible implementation, the processing module 2803 is further configured to:

[0504] The structural control coupling test is deemed passed when the excitation response exceeds the preset amplitude-frequency margin index.

[0505] The structural control coupling test was deemed unsuccessful when the excitation response was less than or equal to the amplitude-frequency margin index.

[0506] In one possible implementation, the amplitude margin is 6 dB or 9 dB.

[0507] In one possible implementation, the sending module 2801 is further configured to:

[0508] Send a test termination command to the flight controller. The test termination command is used to instruct the flight controller to disconnect from at least one actuator.

[0509] In one possible implementation, the frequency range is determined by the characteristics of the aircraft itself and the characteristics of the control system.

[0510] In one possible implementation, the incentive configuration information also includes an incentive type, which is either a single-point incentive or a frequency sweep incentive.

[0511] If the incentive type is frequency sweep incentive, the incentive configuration information also includes the frequency sweep step size and the number of frequency sweeps.

[0512] In one possible implementation, the sending module 2801 is further configured to:

[0513] After completing the test of the aircraft in any configuration, a configuration switching command is sent to the flight control system. The configuration switching command is used to instruct the flight control system to switch the aircraft to other configurations for structural control coupling tests.

[0514] In one possible implementation, the processing module 2803 is further configured to:

[0515] In the open-loop test, if the excitation response under single gain is greater than the preset first amplitude margin index, and the excitation response under double gain is greater than the second amplitude margin index, then the structural control coupling test is deemed to have passed.

[0516] Otherwise, the structural control coupling test is deemed unsuccessful;

[0517] Among them, the first amplitude margin index is greater than the second amplitude margin index.

[0518] The aircraft structure control coupling test device provided in this embodiment can execute the method provided in the above method embodiment. Its implementation principle and technical effect are similar, and will not be described in detail here.

[0519] Figure 29 This is a schematic diagram of a second embodiment of an aircraft structure control coupling test device provided in this application, as shown below. Figure 29 As shown, in Figure 28 Based on the aircraft structure control coupling test device 2800 shown, the aircraft structure control coupling test device 2800 further includes:

[0520] The receiver module 2804 is also used for:

[0521] During the test, the first vibration information of the aircraft's airframe structure was received by the vibration acquisition equipment.

[0522] Receive the second vibration information of the aircraft's airframe structure collected by the inertial navigation system and uploaded by the flight control system during the test;

[0523] The vibration information includes first vibration information and second vibration information.

[0524] The aircraft structure control coupling test device provided in this embodiment can execute the method provided in the above method embodiment. Its implementation principle and technical effect are similar, and will not be described in detail here.

[0525] Figure 30 This is a schematic diagram of the structure of a third embodiment of an aircraft structure control coupling test device provided in this application, as shown below. Figure 30 As shown, the aircraft structure control coupling test device 3000 includes:

[0526] The receiving module 3001 is used to receive the excitation signal generation command sent by the control device;

[0527] Processing module 3002 is used to generate an excitation signal according to the excitation signal generation command and the pre-acquired frequency range, and output the excitation signal to at least one actuator of the aircraft;

[0528] The transmission module 3003 is used to return the vibration information collected by the inertial navigation system of the aircraft to the control equipment.

[0529] In one possible implementation, the receiving module 3001 is further configured to:

[0530] Receive the excitation configuration information sent by the control device, which carries the frequency range.

[0531] In one possible implementation, the processing module 3002 is further configured to:

[0532] The excitation signal is distributed according to a preset distribution matrix to obtain the control quantity for each actuator;

[0533] Each actuator is driven according to the control quantity of each actuator.

[0534] In one possible implementation, the receiving module 3001 is further configured to:

[0535] Receive test status control commands sent by the control equipment. These commands are used to instruct whether to perform an open-loop test or a closed-loop test.

[0536] In one possible implementation, the processing module 3002 is further configured to:

[0537] If the test state control command instructs to perform an open-loop test, then disconnect the channel of the stabilization control signal;

[0538] If the test state control command instructs to perform a closed-loop test, the channel of the stabilization control signal is activated, and a stabilization control signal is generated and superimposed on the excitation signal.

[0539] In one possible implementation, the receiving module 3001 is further configured to:

[0540] Receive parameter adjustment instructions sent by the control equipment.

[0541] In one possible implementation, the processing module 3002 is further configured to:

[0542] According to the parameter adjustment command, the frequency and / or amplitude of the excitation signal are adjusted.

[0543] In one possible implementation, the receiving module 3001 is further configured to:

[0544] Receive the excitation stop command sent by the control equipment.

[0545] In one possible implementation, the processing module 3002 is further configured to:

[0546] The generation of excitation signals is stopped according to the excitation stop instruction.

[0547] In one possible implementation, the receiving module 3001 is further configured to:

[0548] Receive the termination test command sent by the control device.

[0549] In one possible implementation, the processing module 3002 is further configured to:

[0550] Disconnect the flight controller from at least one actuator according to the test termination command.

[0551] In one possible implementation, the incentive configuration information also includes an incentive type, which is either a single-point incentive or a frequency sweep incentive.

[0552] If the incentive type is frequency sweep incentive, the incentive configuration information also includes the frequency sweep step size and the number of frequency sweeps.

[0553] In one possible implementation, the receiving module 3001 is further configured to:

[0554] Receive configuration switching commands sent by the control equipment.

[0555] In one possible implementation, the processing module 3002 is further configured to:

[0556] According to the configuration switching command, the control aircraft is switched to other configurations to conduct structural control coupling tests.

[0557] In one possible implementation, the excitation signal is a single-point signal or a frequency sweep signal when conducting open-loop testing;

[0558] During closed-loop testing, the excitation signal is either a pulse signal or a frequency sweep signal.

[0559] In one possible implementation, the processing module 3002 is further configured to:

[0560] During the closed-loop test, the notch filter is controlled to adjust the stabilization control signal based on the amplitude and phase information of the excitation signal and the stabilization control signal.

[0561] The aircraft structure control coupling test device provided in this embodiment can execute the method provided in the above method embodiment. Its implementation principle and technical effect are similar, and will not be described in detail here.

[0562] Figure 31 This is a schematic diagram of the structure of a computer device provided in this application. Figure 31 As shown, the computer device 3100 provided in this embodiment includes at least one processor 3101, a memory 3102, and an interaction interface 3103. In specific implementation, at least one interaction interface 3103 is used to realize data transmission with the flight control and vibration acquisition equipment in the aircraft, and at least one processor 3101 executes the computer execution instructions stored in the memory 3102, so that at least one processor 3101 executes the above-mentioned aircraft structure control coupling test method.

[0563] The specific implementation process of processor 3101 can be found in the above-mentioned embodiment of the test method for aircraft structure control coupling. Its implementation principle and technical effect are similar, and will not be repeated here.

[0564] Figure 32 This is a schematic diagram of a flight control system provided in this application. Figure 32 As shown, the flight control 102 provided in this embodiment includes at least one processor 3201, a memory 3202, and an interaction interface 3203. In specific implementation, at least one interaction interface 3203 is used to realize data transmission with computer equipment, inertial navigation, and actuators. At least one processor 3201 executes computer execution instructions stored in the memory 3202, so that at least one processor 3201 executes the above-mentioned aircraft structure control coupling test method.

[0565] The specific implementation process of processor 3201 can be found in the above-mentioned embodiment of the test method for aircraft structure control coupling. Its implementation principle and technical effect are similar, and will not be repeated here.

[0566] In the aircraft structure control coupling test provided in this application, to ensure test safety, the propeller needs to be replaced during the test of a rotorcraft or compound wing aircraft. Therefore, this application also provides a test aircraft (i.e., an aircraft after replacing the propeller blades of a rotorcraft or compound wing aircraft), comprising: an aircraft body, Figure 32 The flight controller provided in the illustrated embodiment, and the propeller selected according to the following formula (1) or according to the following formulas (1) and (2):

[0567]

[0568] Where ρ is the atmospheric density, N is the number of propeller blades, π is pi, and C L R is the propeller thrust coefficient, R is the propeller radius, L is the propeller occupancy information, and N is the propeller's thrust coefficient. RPM d is the nominal rotational speed of the propeller. RPM f is the maximum value of the propeller differential speed. min The lowest sweep frequency is I, where I is the inertia of the aircraft under test, and p is the lowest sweep frequency. max M is the maximum angular rate that drives the aircraft to operate, M is the weight of the aircraft under test, and α is a preset proportional threshold.

[0569] For specific replacement and testing procedures, please refer to the description in the method embodiments.

[0570] This application also provides a computer program product, including a computer program that, when executed by a processor, implements the above-described aircraft structure control coupling test method.

[0571] This application also provides a computer-readable storage medium storing computer-executable instructions, which, when executed by a processor, implement the aforementioned aircraft structure control coupling test method.

[0572] Those skilled in the art will understand that all or part of the steps of the above-described method embodiments can be implemented by hardware related to program instructions. The aforementioned program can be stored in a computer-readable storage medium. When executed, the program performs the steps of the above-described method embodiments; and the aforementioned storage medium includes various media capable of storing program code, such as ROM, RAM, magnetic disks, or optical disks.

[0573] Finally, it should be noted that other embodiments of the invention will readily occur to those skilled in the art upon consideration of the specification and practice of the invention disclosed herein. This invention is intended to cover any variations, uses, or adaptations of the invention that follow the general principles of the invention and include common knowledge or customary techniques in the art not disclosed herein, and is not limited to the precise structures described above and shown in the accompanying drawings, and various modifications and changes can be made without departing from its scope. The scope of the invention is limited only by the appended claims.

Claims

1. A method for coupling structural control in an aircraft, characterized in that, include: Send an excitation signal generation command to the flight control system of the aircraft, the excitation signal generation command being used to instruct the flight control system to generate an excitation signal to drive at least one actuator of the aircraft; The vibration information of the aircraft's airframe structure during the test was acquired by the aircraft's inertial navigation and vibration acquisition equipment. The structural control coupling test is determined based on the vibration information.

2. The method according to claim 1, characterized in that, Before sending the excitation signal generation command to the flight control system of the aircraft, the method further includes: The flight controller is sent excitation configuration information, which carries the frequency range used to generate the excitation signal.

3. The method according to claim 2, characterized in that, The method further includes: Send test status control commands to the flight controller, the test status control commands being used to instruct whether to conduct an open-loop test or a closed-loop test.

4. The method according to claim 3, characterized in that, The method further includes: After the flight controller generates a stabilization control signal and superimposes it onto the excitation signal, it determines whether to adjust the frequency and / or amplitude of the excitation signal based on the excitation signal and the stabilization control signal.

5. The method according to any one of claims 1 to 4, characterized in that, The method further includes: If it is determined that the frequency and / or amplitude of the excitation signal needs to be adjusted, a parameter adjustment command is sent to the flight controller, which instructs the flight controller to adjust the frequency and / or amplitude of the excitation signal.

6. The method according to any one of claims 1 to 4, characterized in that, The method further includes: An excitation abort command is sent to the flight controller, the excitation abort command being used to instruct the flight controller to stop generating the excitation signal.

7. The method according to any one of claims 1 to 4, characterized in that, The step of determining whether the structural control coupling test passes based on the vibration information includes: Based on the vibration information, the excitation response of the aircraft during the test is obtained; Based on the excitation response, determine whether the structural control coupling test is successful.

8. The method according to claim 7, characterized in that, The step of determining whether the structural control coupling test passes based on the excitation response includes: The structural control coupling test is deemed passed when the excitation response is greater than the preset amplitude-frequency margin index. The structural control coupling test is deemed to have failed when the excitation response is less than or equal to the amplitude-frequency margin index.

9. The method according to claim 8, characterized in that, The amplitude-frequency margin index is 6dB or 9dB.

10. The method according to any one of claims 1 to 4, characterized in that, The vibration information of the aircraft's airframe structure during the test, acquired by the inertial navigation and vibration acquisition equipment, includes: During the test, the first vibration information of the aircraft's airframe structure acquired by the vibration acquisition device was received; Receive the second vibration information of the aircraft's airframe structure during the test, which is collected by the inertial navigation system and uploaded by the flight control system; The vibration information includes the first vibration information and the second vibration information.

11. The method according to any one of claims 1 to 4, characterized in that, The method further includes: A test termination command is sent to the flight controller, the test termination command being used to instruct the flight controller to disconnect from at least one actuator.

12. The method according to any one of claims 2 to 4, characterized in that, The frequency range is determined by the characteristics of the aircraft itself and the characteristics of the control system.

13. The method according to any one of claims 2 to 4, characterized in that, The incentive configuration information also includes an incentive type, which is either a single-point incentive or a frequency sweep incentive. If the incentive type is frequency sweep incentive, the incentive configuration information also includes frequency sweep step size and frequency sweep count.

14. The method according to any one of claims 1 to 4, characterized in that, The method further includes: After completing the test of the aircraft in any configuration, a configuration switching command is sent to the flight controller. The configuration switching command is used to instruct the flight controller to control the aircraft to switch to other configurations for structural control coupling test.

15. The method according to claim 7, characterized in that, The step of determining whether the structural control coupling test passes based on the excitation response includes: In the open-loop test, if the excitation response under single gain is greater than the preset first amplitude margin index, and the excitation response under double gain is greater than the second amplitude margin index, then the structural control coupling test is deemed to have passed. Otherwise, the structural control coupling test is deemed unsuccessful; The first amplitude margin index is greater than the second amplitude margin index.

16. A method for coupling structural control in an aircraft, characterized in that, The method, applied to flight control of aircraft, includes: Receive excitation signal generation instructions sent by the control equipment; An excitation signal is generated according to the excitation signal generation command and the pre-acquired frequency range, and the excitation signal is output to at least one actuator of the aircraft; The vibration information collected by the inertial navigation system of the aircraft is returned to the control device.

17. The method according to claim 16, characterized in that, Before the excitation signal sent by the receiving control device generates a command, the method further includes: The system receives excitation configuration information sent by the control device, wherein the excitation configuration information carries the frequency range.

18. The method according to claim 16, characterized in that, The step of outputting the excitation signal to at least one actuator of the aircraft includes: The excitation signal is distributed according to a preset distribution matrix to obtain the control quantity for each actuator; Each actuator is driven according to the control quantity of each actuator.

19. The method according to any one of claims 16 to 18, characterized in that, The method further includes: Receive test status control instructions sent by the control device, the test status control instructions being used to indicate whether to perform an open-loop test or a closed-loop test; If the test state control command instructs to perform an open-loop test, then the channel of the stabilization control signal is disconnected; If the test state control command instructs to perform a closed-loop test, the channel of the stabilization control signal is activated, and a stabilization control signal is generated and superimposed on the excitation signal.

20. The method according to any one of claims 16 to 18, characterized in that, The method further includes: Receive parameter adjustment instructions sent by the control device; The frequency and / or amplitude of the excitation signal are adjusted according to the parameter adjustment command.

21. The method according to any one of claims 16 to 18, characterized in that, The method further includes: Receive the excitation stop command sent by the control device; The generation of the excitation signal is stopped according to the excitation stop instruction.

22. The method according to any one of claims 16 to 18, characterized in that, The method further includes: Receive the termination test command sent by the control device; According to the termination test command, disconnect the connection between the flight controller and the at least one actuator.

23. The method according to any one of claims 16 to 18, characterized in that, The incentive configuration information also includes an incentive type, which is either a single-point incentive or a frequency sweep incentive. If the incentive type is frequency sweep incentive, the incentive configuration information also includes frequency sweep step size and frequency sweep count.

24. The method according to any one of claims 16 to 18, characterized in that, The method further includes: Receive configuration switching instructions sent by the control device; According to the configuration switching command, the aircraft is controlled to switch to other configurations to conduct structural control coupling tests.

25. The method according to any one of claims 16 to 18, characterized in that, When conducting open-loop tests, the excitation signal can be a single-point signal or a frequency sweep signal; During closed-loop testing, the excitation signal is either a pulse signal or a frequency sweep signal.

26. The method according to any one of claims 16 to 18, characterized in that, During the closed-loop test, the notch filter is controlled to adjust the stabilization control signal based on the amplitude and phase information of the excitation signal and the stabilization control signal.

27. A test device for coupling control of an aircraft structure, characterized in that, include: The transmitting module is used to send an excitation signal generation command to the flight control system of the aircraft. The excitation signal generation command is used to instruct the flight control system to generate an excitation signal to drive at least one actuator of the aircraft. The acquisition module is used to acquire vibration information of the aircraft's airframe structure during the test, obtained by the aircraft's inertial navigation and vibration acquisition equipment. The processing module is used to determine whether the structural control coupling test has passed based on the vibration information.

28. A test device for coupling control of an aircraft structure, characterized in that, include: The receiving module is used to receive excitation signal generation instructions sent by the control equipment; The processing module is configured to generate an excitation signal based on the excitation signal generation command and a pre-acquired frequency range, and output the excitation signal to at least one actuator of the aircraft. The transmission module is used to return the vibration information collected by the inertial navigation system of the aircraft to the control device.

29. A computer device, characterized in that, include: Memory, processor, and interface; The memory stores computer-executed instructions; The processor executes computer execution instructions stored in the memory, causing the computer device to perform the aircraft structure control coupling test method as described in any one of claims 1 to 15.

30. A flight control system, characterized in that, include: Memory, processor, and interface; The memory stores computer-executed instructions; The processor executes computer execution instructions stored in the memory, causing the computer device to perform the aircraft structure control coupling test method as described in any one of claims 16 to 26.

31. An experimental aircraft, characterized in that, include: The aircraft body, the flight control system as described in claim 30, and the propeller selected according to formula (1) or formulas (1) and (2) as follows: Where ρ is the atmospheric density, N is the number of propeller blades, π is pi, and C L R is the propeller thrust coefficient, R is the propeller radius, L is the propeller occupancy information, and N is the propeller's thrust coefficient. RPM d is the nominal rotational speed of the propeller. RPM f is the maximum value of the propeller differential speed. min The lowest sweep frequency is I, where I is the inertia of the aircraft under test, and p is the lowest sweep frequency. max M is the maximum angular rate that drives the aircraft to operate, M is the weight of the aircraft under test, and α is a preset proportional threshold.

32. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores computer-executable instructions, which, when executed by a processor, are used to implement the aircraft structure control coupling test method as described in any one of claims 1 to 26.

33. A computer program product, characterized in that, It includes a computer program that, when executed by a processor, implements the aircraft structure control coupling test method according to any one of claims 1 to 26.

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