Aircraft structure control coupling test system

Abstract

This application provides an aircraft structure control coupling test system, relating to the field of aircraft. Specifically, the system includes a control device and a vibration acquisition device mounted on the aircraft's airframe structure. The control device is connected to the aircraft's flight controller and is responsible for configuring the required frequency sweep range for the test. The flight controller generates an excitation signal based on the set frequency sweep range and transmits it to at least one actuator to excite the aircraft's airframe structure. The aircraft's inertial navigation system and vibration acquisition device work together to collect vibration information of the aircraft's airframe structure during the execution process. The control device analyzes the collected vibration information and determines whether the coupling test passes. This design improves the versatility of the test system, enabling it to adapt to the needs of different types and scales of aircraft structure control coupling tests.

Description

Technical Field

[0001] This application relates to the field of aircraft, and more particularly to an aircraft structure control coupling test system. 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. Before bringing such aircraft to market, structural control coupling tests are usually required to solve the coupling problem between its structure and control system.

[0003] Currently, structural control coupling tests typically require the use of multiple independent instruments. For example, a frequency sweep generator or signal generator is used as an excitation source to generate signals of different frequencies to stimulate the structural response of the aircraft. Meanwhile, equipment such as oscilloscopes are used to observe and record the test results in real time.

[0004] However, existing structural control coupling test setups are complex, involving the coordinated use of multiple independent devices, which increases the difficulty and cost of testing. Furthermore, the aforementioned excitation source application methods are difficult to implement on eVTOL systems. Therefore, there is an urgent need to develop a more universal aircraft structural control coupling test system to improve the applicability of the tests. Summary of the Invention

[0005] This application provides an aircraft structure control coupling test system to improve the applicability of the test.

[0006] In a first aspect, embodiments of this application provide an aircraft structure control coupling test system, comprising:

[0007] Control equipment, and vibration acquisition equipment installed on the aircraft's airframe structure;

[0008] The control device is connected to the flight controller of the aircraft and is used to configure the frequency range corresponding to the test in the flight controller. The flight controller is used to generate an excitation signal according to the frequency range and output it to at least one actuator.

[0009] The inertial navigation system and the vibration acquisition device of the aircraft are used to collect vibration information on the airframe structure of the aircraft during execution;

[0010] The control device is also used to determine whether the structural control coupling test has passed based on the vibration information.

[0011] In one possible implementation, the control device is further configured to configure at least one of the generation, termination, frequency, and amplitude of the excitation signal.

[0012] In one possible implementation, the control device is specifically used to determine the excitation response based on the vibration information, and to determine whether the structural control coupling test passes based on the excitation response.

[0013] In one possible implementation, the control device is specifically used to: determine that the structural control coupling test has passed when the excitation response is greater than a preset amplitude-frequency margin index, and determine that the structural control coupling test has failed when the excitation response is less than or equal to the amplitude-frequency margin index.

[0014] In one possible implementation, the amplitude margin index is 6dB or 9dB.

[0015] In one possible implementation, a first switch is provided between the flight controller and the at least one actuator;

[0016] The control device is also used to control the on / off state of the first switch.

[0017] In one possible implementation, the flight controller is also used to distribute the excitation signal to each actuator according to a preset allocation matrix.

[0018] In one possible implementation, the flight controller is further configured to generate a stabilization control signal based on the excitation response and superimpose it onto the excitation signal during a closed-loop test.

[0019] In one possible implementation, a second switch is provided on the output path of the excitation signal. The second switch is located before the excitation signal is superimposed with the stabilization control signal. The second switch is used to control the access or disconnection of the excitation signal.

[0020] In one possible implementation, a third switch is provided on the output path of the stabilization control signal. The third switch is located before the excitation signal and the stabilization control signal are superimposed. The third switch is used to control the state of the experiment to be either open-loop or closed-loop.

[0021] In one possible implementation, the number of vibration acquisition devices is multiple;

[0022] Multiple vibration acquisition devices are respectively installed at the inertial navigation system, each actuator, and at least one location representing the vibration mode of the structure.

[0023] In one possible implementation, the location of the at least one characterizing structural vibration includes: the wingtip and / or tailtip of the aircraft.

[0024] In one possible implementation, the aircraft is a fixed-wing aircraft, and the at least one actuator includes: servo motors respectively disposed at the elevator, rudder and aileron positions of the fixed-wing aircraft, or servo motors respectively disposed at the elevator rudder and aileron positions of the fixed-wing aircraft.

[0025] In one possible implementation, the aircraft is a rotorcraft, and the at least one actuator includes a motor that drives the propeller.

[0026] In one possible implementation, the aircraft is a compound wing aircraft, and the at least one actuator includes:

[0027] Servo mechanisms located at the elevator, rudder, and aileron positions, and motors driving the propeller; or,

[0028] The servo motors are located at the positions of the elevator and ailerons, as well as the motors that drive the propeller.

[0029] In one possible implementation, the system further includes a support device for supporting the aircraft in different configurations during testing.

[0030] In one possible implementation, the support device includes a flexible suspension device or a support frame;

[0031] The elastic suspension device is used to suspend the aircraft for open-loop and closed-loop tests of takeoff and landing configurations, transition configurations, and cruise configurations.

[0032] The support frame is used to support the aircraft in open-loop and closed-loop tests of takeoff and landing configurations.

[0033] In one possible implementation, the frequency of the support device is lower than a preset frequency threshold for the first-order structural mode frequency of the aircraft.

[0034] In one possible implementation, the frequency domain threshold is one-third of the frequency of the first-order structural mode.

[0035] In one possible implementation, the system further includes: a control surface elastic tooling and / or a counterweight for increasing the excitation force of the control surface during frequency sweep;

[0036] The control surface elastic tooling is mounted on the control surface of the aircraft;

[0037] The counterweight is located on the rear axle of the aircraft's control surfaces.

[0038] In one possible implementation, the control device is also used to control the flight controller to drive the configuration switching of the aircraft.

[0039] In one possible implementation, when the aircraft is a rotorcraft or a compound wing aircraft,

[0040] The propeller of the rotorcraft or compound wing aircraft is selected according to the following formula (1) or according to the following formulas (1) and (2):

[0041]

[0042]

[0043] 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 differential speed of the propeller. 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.

[0044] In one possible implementation, the control device is specifically used for:

[0045] 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.

[0046] Otherwise, the structural control coupling test is deemed unsuccessful;

[0047] The first amplitude margin index is greater than the second amplitude margin index.

[0048] The aircraft structural control coupling test system provided in this application relates to the field of aircraft. Specifically, the system includes control equipment and vibration acquisition equipment mounted on the aircraft's airframe structure. The control equipment is connected to the aircraft's flight control system, enabling flexible configuration of the required frequency sweep range for the test. This allows for the generation and transmission of precise excitation signals to at least one actuator, achieving effective excitation of the aircraft's airframe structure. Simultaneously, the combined use of the aircraft's inertial navigation system and vibration acquisition equipment ensures comprehensive capture and real-time monitoring of vibration information from the aircraft's airframe structure. Furthermore, the control equipment analyzes the acquired vibration information to determine the results of the structural control coupling test. This design not only simplifies the testing process and improves the accuracy and reliability of the data but also significantly enhances the versatility of the test system, enabling it to adapt to the needs of structural control coupling tests for different types and scales of aircraft. Attached Figure Description

[0049] 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.

[0050] Figure 1 A schematic diagram of the scenario for the aircraft structure control coupling test system provided in this application;

[0051] Figure 2 A schematic diagram of an embodiment of the aircraft structure control coupling test system provided in this application;

[0052] Figure 3 A schematic diagram of Embodiment 2 of the aircraft structure control coupling test system provided in this application;

[0053] Figure 4 A schematic diagram of Embodiment 3 of the aircraft structure control coupling test system provided in this application;

[0054] Figure 5 A schematic diagram of Embodiment 4 of the aircraft structure control coupling test system provided in this application;

[0055] Figure 6 This is a schematic diagram of Embodiment 5 of the aircraft structure control coupling test system provided in this application.

[0056] Explanation of reference numerals in the attached figures:

[0057] 101-Control equipment; 102-Vibration acquisition equipment; 103-Aircraft; 1031-Flight control; 1032-Inertial navigation; 1033-Actuator; 200-Aircraft structure control coupling test system; 300-Aircraft structure; 301-Main control PC.

[0058] The accompanying drawings have illustrated 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 specific embodiments. Detailed Implementation

[0059] 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.

[0060] First, let me explain the terms used in this application:

[0061] 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;

[0062] 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;

[0063] 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;

[0064] 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.

[0065] Flight control: is an electronic system on an aircraft used to automatically or manually control the aircraft's flight attitude and trajectory;

[0066] 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.

[0067] In recent years, with the rapid development of materials science, energy technology, and control technology, the feasibility of eVTOL research and development has been fully verified. 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.

[0068] Currently, in conducting structural control coupling tests on aircraft, existing technologies typically rely on multiple independent instruments and devices. A sweep generator or signal generator is used as the excitation source, producing a tunable frequency signal to stimulate the aircraft's structure. Oscilloscopes and data loggers are used to capture and record the dynamic response of the structure. These devices, through precise configuration and coordination, enable synchronous data acquisition and analysis. The data acquisition system and specialized software tools further perform modal and spectral analysis on this data to identify the structure's inherent frequencies and modal shapes, thereby helping test personnel understand the coupling characteristics between the structure and the control system.

[0069] However, existing structural control coupling test systems suffer from complex configurations and cumbersome operations, involving the coordinated use of multiple independent devices, which increases the difficulty and cost of testing. Furthermore, traditional excitation source application methods are difficult to implement on eVTOL aircraft, failing to effectively simulate the unique coupling characteristics of their structure and control systems. Therefore, there is an urgent need to develop a more universal aircraft structural control coupling test system to improve the applicability of the tests.

[0070] To address the aforementioned technical problems, the inventors, during their research on aircraft structure control coupling test systems, discovered that eVTOL aircraft primarily employ rotor or compound wing designs. However, existing structure control coupling test systems are mainly designed for fixed-wing aircraft, typically implemented by adding excitation sources to control surfaces. This method is not effectively applicable to eVTOL aircraft. Furthermore, existing systems are relatively complex, usually requiring independent frequency sweepers or signal generators as excitation sources, and relying on equipment such as oscilloscopes to observe test results. Therefore, the inventors considered whether existing aircraft structure control coupling test systems could be improved and optimized to solve the problem of not being able to directly add excitation sources to eVTOL aircraft using traditional methods. Specifically, this application provides an aircraft structure control coupling test system, which includes control equipment and vibration acquisition equipment mounted on the aircraft's airframe structure. The control equipment is connected to the aircraft's flight controller and is used to configure the required frequency range for the test within the flight controller. The flight controller generates an excitation signal based on this frequency range and outputs it to at least one actuator. The aircraft's inertial navigation and vibration acquisition equipment are used to collect vibration information from the aircraft's airframe structure during execution. The control equipment uses this vibration information to determine whether the structural control coupling test has passed. This system significantly improves the versatility of structural control coupling tests by effectively implementing excitation sources on eVTOL aircraft and integrating vibration acquisition and analysis functions. It can adapt to different types of aircraft structures and control requirements, making structural control coupling tests not only more reliable on eVTOL aircraft but also widely applicable to other new types of aircraft, thus overcoming the limitations of existing technologies.

[0071] Figure 1A schematic diagram of the scenario for the aircraft structure control coupling test system provided in this application. Figure 1 As shown, the scenario of this aircraft structure control coupling test system includes control equipment 101, vibration acquisition equipment 102, and aircraft 103. Although Figure 1 Only one control device 101, one vibration acquisition device 102, and one aircraft 103 are shown in the diagram, but it should be understood that there may be two or more control devices 101, vibration acquisition devices 102, and aircraft 103.

[0072] The control device 101 is primarily used to configure the frequency range corresponding to the test in the flight control system and to determine whether the structural control coupling test has passed based on vibration information. It is worth noting that the control device 101 can be a single terminal device, such as a laptop or desktop computer. The control device 101 can also 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.

[0073] The vibration acquisition device 102 is mainly used to collect vibration information from the airframe structure of the aircraft 103 during execution. It is worth noting that the vibration acquisition device 102 can be a single sensor device, such as an accelerometer. The vibration acquisition device 102 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 102. It is worth noting that the airframe structure refers to the fuselage of the aircraft, which is a component of the overall aircraft structure.

[0074] The overall structure and various systems of aircraft 103 are the objects of structural control coupling tests. Aircraft 103 includes, but is not limited to, rotorcraft, compound wingcraft, or fixed-wingcraft. It is worth noting that aircraft 103 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 aircraft 103, in order to adapt to different testing needs and application scenarios.

[0075] Specifically, the aircraft 103 is equipped with a flight control system, inertial navigation system, and actuators. The control device 101 is connected to the flight control system of the aircraft 103 and is used to configure the frequency range required for the test. The flight control system generates excitation signals according to this frequency range and outputs them to at least one actuator. The inertial navigation system of the aircraft 103 and the vibration acquisition device 102 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.

[0076] 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.

[0077] 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.

[0078] Figure 2 This is a schematic diagram of an embodiment of the aircraft structure control coupling test system provided in this application. Figure 2 As shown, the aircraft structure control coupling test system 200 includes a control device 101 and a vibration acquisition device 102 mounted on the airframe structure of the aircraft 103. The aircraft 103 is equipped with a flight controller 1031, an inertial navigation system 1032, and an actuator 1033.

[0079] The function and type of each device in the aircraft structure control coupling test system 200 will be described in detail below.

[0080] Specifically, in the aircraft structure control coupling test system 200, the control device 101 is mainly used to configure the frequency range corresponding to the test in the flight control 1031 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.

[0081] Vibration acquisition device 102 is mounted on the fuselage structure of aircraft 103 to collect vibration information from the fuselage structure of aircraft 103 during the test. Vibration acquisition device 102 includes, but is not limited to, single sensor devices or combinations of multiple sensor devices; this application does not specifically limit its use. It is worth noting that the vibration acquisition device installed during the test should be able to acquire the following information: it should be able to reflect the vibration along the rotor thrust axis at each propeller motor; it should be able to reflect the vibration along the deflection direction at each control surface rear axle; it should be able to characterize the main mode shapes within 10 times the control bandwidth frequency; and it should be able to reflect the three-axis acceleration at the main and backup inertial navigation systems. The vibration acquisition device is an accelerometer used to acquire acceleration data.

[0082] The overall structure and various systems of aircraft 103 are the objects of structural control coupling tests. Aircraft 103 includes, but is not limited to, rotorcraft, compound-wing aircraft, or fixed-wing aircraft; this application does not specifically limit this. The overall structure refers to the physical structure of the entire aircraft, including the airframe structure and other auxiliary structural parts. Each system refers to the various subsystems within the aircraft related to structural performance, including but not limited to flight control 1031, navigation system, communication system, and actuation system.

[0083] It should be noted that the aircraft involved in this application mainly refers to powered lift aircraft.

[0084] The flight controller 1031 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 1033. The flight controller 1031 includes, but is not limited to, embedded controllers, other computing platforms applicable to aircraft control, etc., which are not specifically limited in this application.

[0085] The inertial navigation system 1032 is used to collect vibration information from the airframe structure of the aircraft 103 during the test, and together with the vibration acquisition device 102, it provides feedback data. The inertial navigation system 1032 includes a primary inertial navigation system and a backup inertial navigation system. The data recording frequency of the primary and backup inertial navigation systems should be no less than 5 times the control bandwidth, and it is mainly used to obtain linear acceleration and angular velocity.

[0086] The actuator 1033 performs corresponding actions based on the excitation signal generated by the flight controller 1031 to influence the structural dynamics of the aircraft 103, thereby achieving the objective of the structural control coupling test. It is worth noting that different types of aircraft have different actuators 1033, 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.

[0087] The connection methods between the various devices in the aircraft structure control coupling test system 200 will be described below.

[0088] The control device 101 is connected to the flight controller 1031 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.

[0089] The flight controller 1031 is connected to the actuator 1033 via a signal line or a dedicated control bus to send excitation signals and control its actions.

[0090] The vibration acquisition device 102 is connected to the actuator 1033 via a data bus or wireless connection to acquire vibration information of the aircraft 103's airframe structure in real time. Simultaneously, the vibration acquisition device 102 is connected to the control device 101 via a data bus or wireless connection to transmit the acquired vibration information in real time.

[0091] The inertial navigation system 1032 is mounted on the airframe structure of the aircraft 103 and is connected to the flight controller 1031 via signal lines or a dedicated control bus to provide real-time vibration information.

[0092] The above connection method can ensure that data and commands can be transmitted efficiently and accurately between the devices within the aircraft structure control coupling test system 200, thereby supporting the smooth progress of the structure control coupling test.

[0093] It should be noted that before the test begins, the client should provide the testing unit with the necessary technical support documents, mainly including 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.

[0094] The execution logic of the aircraft structure control coupling test system 200 will be described in detail below.

[0095] 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 1031 via a wired or wireless communication interface. After receiving the frequency range setting information, the flight controller 1031 generates the corresponding excitation signal and sends it to the actuator 1033 via a signal line or control bus.

[0096] The actuator 1033 performs a specific action based on the received excitation signal to influence the structural dynamics of the aircraft 103.

[0097] Meanwhile, vibration acquisition device 102 and inertial navigation system 1032 are respectively installed on the airframe structure of aircraft 103 to collect vibration information in real time. Among them, vibration acquisition device 102 transmits vibration information to control device 101 through data bus or wireless connection, and inertial navigation system 1032 transmits vibration information to flight controller 1031 through signal line or control bus.

[0098] Furthermore, the flight controller 1031 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 102 and the vibration information sent by the inertial navigation system 1032 through the flight controller 1031.

[0099] The frequency range refers to the interval between the lowest and highest frequencies of the excitation signal applied in the structural control coupling test of an aircraft, and is usually determined by the ground vibration test (GVT). In the structural control coupling test of an aircraft, by setting the frequency range appropriately, resonant frequencies can be identified and avoided, thereby ensuring the stability and reliability of the aircraft under various dynamic conditions.

[0100] In one possible implementation, the frequency range is determined by the characteristics of the aircraft itself and the characteristics of the control system.

[0101] 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.

[0102] It is worth noting that the aircraft's inherent characteristics are primarily determined based on the results of ground vibration tests (GVT). 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 test and the accuracy of the data.

[0103] 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.

[0104] 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.

[0105] The excitation signal refers to the input signal applied by the flight controller 1031 to the actuator 1033 in the aircraft structure control coupling test, with the purpose of inducing a dynamic response of the aircraft structure. It is worth noting that the excitation signal is used to simulate the dynamic conditions encountered by the aircraft 103 in actual operation. By applying the excitation signal, the response characteristics of the aircraft structure can be tested, potential resonant frequencies can be identified, and the stability and reliability of the structure can be evaluated.

[0106] 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.

[0107] 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.

[0108] 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 1033 via the signal lines or control bus of the flight controller 1031. Finally, upon receiving the excitation signal, the actuator 1033 performs corresponding physical actions to influence the structural dynamics of the aircraft 103.

[0109] When the actuator 1033 performs the corresponding physical action, the vibration acquisition device 102 and the inertial navigation system 1032 respectively acquire vibration information of the aircraft 103's airframe structure. The vibration information refers to the dynamic response data of the aircraft's airframe structure acquired during the test. It is worth noting that because the vibration acquisition device 102 and the inertial navigation system 1032 are located at different positions on the aircraft 103, the acquired vibration information differs; the vibration acquisition device 102 provides the local structural response, while the inertial navigation system 1032 provides the global motion characteristics.

[0110] Specifically, vibration acquisition device 102 primarily focuses on the dynamic response of the aircraft's airframe structure, acquiring vibration information in the form of acceleration data to identify the resonant frequencies of the airframe structure. Inertial navigation system 1032 focuses on the overall motion characteristics of the aircraft 103, acquiring vibration information including linear acceleration data and angular velocity to provide analysis of the motion and dynamic behavior of the aircraft 103. The combination of these two vibration information allows control device 101 to comprehensively evaluate the dynamic performance and structural stability of the aircraft.

[0111] In one possible implementation, the vibration acquisition device 102 is installed on critical structural parts of the aircraft 103, such as the wings, fuselage, or other areas significantly affected by vibration, to directly measure the dynamic response and vibration characteristics of these parts. It is worth noting that the specific installation location of the vibration acquisition device 102 is determined according to the requirements of the ground vibration test, but the locations of the inertial navigation system 1032 and the motors must be where the vibration acquisition device 102 is installed.

[0112] The inertial navigation system 1032 is mounted near the center of gravity of the aircraft 103 or another suitable location to ensure accurate measurement of the linear acceleration and angular velocity of the entire aircraft. The inertial navigation system 1032 primarily provides global motion characteristic data to aid in the analysis of the overall dynamic behavior and motion of the aircraft 103.

[0113] It should be noted that before the aircraft structure control coupling test begins, communication tests between the flight controller 1031, inertial navigation system 1032, motors, electronic control, vibration acquisition equipment 102, and control equipment 101 need to be completed.

[0114] It is worth noting that the aircraft structure control coupling test system 200 is mainly used for conducting aircraft structure control coupling tests. The primary purpose of these tests is to obtain the aircraft structure control coupling characteristics and determine the stability of the structure control coupling loop by conducting servo-elasticity tests on the ground. Furthermore, the test results can be used to correct the mathematical model in aerodynamic servo-elasticity calculations, making the aerodynamic servo-elasticity analysis more reliable. In addition, the aircraft structure control coupling test is a verification test, conducted after the ground vibration test, focusing on excitation conditions near the structural modal frequencies.

[0115] The aircraft structure control coupling test system provided in this application significantly improves the universality of the test and reduces the complexity of the test process by integrating control equipment and vibration acquisition equipment. Unlike existing systems that rely on multiple independent devices such as sweep frequency meters, signal generators, and oscilloscopes, this system reduces dependence on external equipment by integrating these functions into a single control device. The control device can directly configure the frequency range and control the flight controller to generate corresponding excitation signals. This method of generating excitation sources is applicable to all types of aircraft, not just fixed-wing aircraft. Through this integrated design, the system not only simplifies the operation process and reduces test costs, but also improves the accuracy and synchronization of data acquisition, making it more flexible in adapting to different types of aircraft, especially meeting the research and development and testing needs of eVTOL aircraft.

[0116] In one possible implementation, the control device 101 is further configured to configure at least one of the following: the generation, termination, cancellation, frequency, and amplitude of the excitation signal.

[0117] Specifically, the occurrence of the excitation signal refers to the moment when the excitation signal is first applied to the airframe structure of aircraft 103, marking the start of the test. Controlling the occurrence of the excitation signal allows test personnel to begin applying vibration or other forms of excitation at specific time points to observe the initial dynamic response of the airframe structure of aircraft 103.

[0118] The termination of the excitation signal refers to the moment when the excitation signal is no longer applied to the structure of aircraft 103, marking a phased pause in the test. Controlling the termination of the excitation signal helps test personnel to stop the excitation at the appropriate time to avoid excessive vibration that could damage the structure.

[0119] The termination of the excitation signal refers to the moment when the excitation signal is completely stopped being applied to the structure of aircraft 103, marking the end of the test or the completion of a phase. Unlike the cessation of the excitation signal, the termination of the excitation signal not only means the cessation of the application of the excitation signal, but also represents the end of the entire test phase or the transition to the next phase.

[0120] The excitation signal frequency can be a specific single-point frequency or a frequency range. Testers can choose to set a fixed frequency or define a frequency range, which includes the start frequency, end frequency, sweep step size, and number of sweeps. If a single-point frequency is set, the flight controller 1031 will generate an excitation signal at that specific frequency to test the airframe structure's response characteristics at that frequency. If a frequency range is set, the flight controller 1031 will scan and generate excitation signals of different frequencies within that range.

[0121] The amplitude of the excitation signal can be a fixed value or a range. If a fixed amplitude is set, the flight controller 1031 generates an excitation signal with that specific intensity to test the airframe structure's response under that amplitude condition. If an amplitude range is set, the flight controller 1031 can adjust the signal intensity within that range to simulate different amplitude conditions.

[0122] By precisely configuring the generation, termination, frequency, and amplitude of the excitation signal, the test system can more comprehensively simulate actual operating conditions and provide more reliable test data.

[0123] 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, but the test does not end. This allows the tester to pause the application of the excitation signal during the test for intermediate data collection or adjustment. When the set termination time is reached, the control device 101 completely stops the output of the excitation signal, marking the end of the test phase. In this way, the tester can flexibly control the termination and interruption of the signal, ensuring that each test phase is completed according to the predetermined requirements.

[0124] Meanwhile, the software interface of the control device 101 allows the test personnel to set the frequency and amplitude of the excitation signal. For frequency setting, the test personnel 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 will transmit these setting parameters to the flight controller 1031, which will generate the corresponding excitation signal according to the settings. If a frequency range is set, the flight controller 1031 will automatically scan and generate excitation signals of different frequencies within that range to identify the natural frequencies and resonance phenomena of the aircraft structure.

[0125] Similarly, for amplitude settings, the tester can choose to input a fixed amplitude value or an amplitude range. The control device 101 will transmit these settings to the flight controller 1031, which will then generate an excitation signal with the corresponding intensity.

[0126] 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.

[0127] The aircraft structure control coupling test system provided in this embodiment has parameter configuration capabilities. Specifically, the control equipment can precisely configure at least one parameter among the generation, termination, cancellation, frequency, and amplitude of the excitation signal. This design significantly improves the flexibility and data accuracy of the test system, enabling it to adapt to various aircraft types and their complex testing requirements. Through precise configuration of multiple parameters, the test system can perform various types of tests. With these precise parameter configurations, the test system can not only provide high-quality test data, supporting more effective design optimization and control system adjustments, but also simplify operating procedures, reduce test complexity and cost, thereby significantly improving test efficiency and reliability.

[0128] In one possible implementation, the control device 101 is specifically used to determine the excitation response based on the vibration information, and to determine whether the structural control coupling test has passed based on the excitation response.

[0129] Specifically, during the test, the control device 101 determines the excitation response based on the vibration information sent by the vibration acquisition device 102 and the vibration information sent by the inertial navigation system 1032 through the flight control system 1031, and determines whether the structural control coupling test is successful based on the excitation response.

[0130] Excitation response refers to the dynamic reaction exhibited by the aircraft's structure when subjected to an excitation signal.

[0131] Optionally, in one specific implementation, the control device 101 first preprocesses the vibration information provided by the vibration acquisition device 102 and the inertial navigation system 1032 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. These parameters form the basis of the excitation response, reflecting the dynamic behavior of the structure under the excitation signal. Finally, these parameters are compared with a set standard range to evaluate whether the dynamic performance of the structure meets the design requirements. If the excitation response parameters measured under all test conditions are consistent with the set standard range and within the allowable error range, the structural control coupling test can be considered successful, indicating that the structural design of the aircraft 103 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 103.

[0132] The aircraft structure control coupling test system provided in this embodiment can effectively determine the excitation response characteristics of the aircraft under the action of excitation signals through precise analysis and processing of vibration information. This not only improves the accuracy and reliability of test results, but also identifies potential problems in structural design in a timely manner, thereby reducing the number of design iterations and costs.

[0133] In one possible implementation, the control device 101 can determine whether the structural control coupling test is successful based on the excitation response and a preset amplitude-frequency margin index. Specifically, the structural control coupling test is determined to have passed when the excitation response is greater than the preset amplitude-frequency margin index, and the structural control coupling test is determined to have failed when the excitation response is less than or equal to the amplitude-frequency margin index.

[0134] The preset amplitude-frequency margin index is a standard used to measure the dynamic performance of an aircraft's airframe structure. The preset amplitude-frequency margin index includes the vibration amplitude and frequency response characteristics of the aircraft's airframe structure at different frequencies. Specifically, the amplitude-frequency margin index involves amplitude margin and frequency margin. Amplitude margin refers to the maximum vibration amplitude that the airframe structure can withstand at a specific frequency without leading to performance degradation or instability. Frequency margin refers to the response characteristics of the airframe structure at different frequencies to ensure that the airframe structure maintains stability at these frequencies.

[0135] Specifically, when the excitation response under all test conditions is greater than the preset amplitude-frequency margin, it indicates that the aircraft airframe structure exhibits good stability and sufficient performance margin under dynamic excitation. This means that the structural design can withstand dynamic stress and vibration under actual flight conditions, and therefore the structural control coupling test is considered passed. Conversely, when the excitation response under any test condition is less than or equal to the amplitude-frequency margin, it indicates that the aircraft airframe structure exhibits instability or insufficient performance margin under dynamic excitation. This will cause the airframe structure to be unable to effectively cope with dynamic stress and vibration under actual flight conditions, thus affecting the safety and performance of the aircraft. Therefore, in this case, the structural control coupling test is considered failed.

[0136] 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 indicates 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.

[0137] It is worth noting that the amplitude margin is 6dB or 9dB, depending on the stage of the aircraft's lifespan and its required performance.

[0138] In one specific embodiment, the control device 101 is specifically used for:

[0139] 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.

[0140] Otherwise, the structural control coupling test is deemed unsuccessful;

[0141] Among them, the first amplitude margin index is greater than the second amplitude margin index.

[0142] Specifically, the excitation response refers to the swept frequency response at all structural modal frequencies.

[0143] 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 exceeds a preset first amplitude margin index, and the excitation response under double-gain conditions exceeds a second amplitude margin index, then the structural control coupling test is deemed to have passed; otherwise, the structural control coupling test is deemed to have failed. In this case, optionally, the first amplitude margin index is 9dB, and the second amplitude margin index is 0dB.

[0144] 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, optionally, the first amplitude margin index is 6dB, and the second amplitude margin index is 0dB.

[0145] 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.

[0146] In one 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.

[0147] 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.

[0148] 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.

[0149] 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.

[0150] The aircraft structure control coupling test system provided in this embodiment 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.

[0151] In one possible implementation, Figure 2 Based on the aircraft structure control coupling test system 200 shown, a first switch is provided between the flight controller 1031 and at least one actuator 1033, and the control device 101 is also used to control the on and off of the first switch.

[0152] Specifically, the primary function of the first switch is to control the on / off state of the signal. It is managed by control device 101 to ensure smooth signal transmission under normal operating conditions. However, in an emergency, control device 101 can quickly disconnect the first switch to protect the aircraft's structure from potential damage caused by overload or other abnormal conditions.

[0153] The aircraft structural control coupling test system provided in this embodiment achieves multiple objectives by setting a first switch between the flight controller and at least one actuator, and utilizing control equipment to realize the on / off function of this switch. First, under normal operating conditions, the system can efficiently conduct structural control coupling tests, ensuring smooth signal transmission and stable system operation. Second, in emergency situations, the control equipment can quickly disconnect the first switch, thereby providing additional protection for the aircraft. This design not only prevents damage that may be caused by overload or other abnormal conditions, but also adds a layer of protection for the safe operation of the aircraft.

[0154] In one possible implementation, the flight controller 1031 in the aircraft structure control coupling test system 200 is also used to distribute the excitation signal to each actuator 1033 according to a preset distribution matrix.

[0155] Specifically, after the flight controller 1031 generates an excitation signal, it distributes it to each actuator 1033 according to the preset distribution matrix.

[0156] The preset allocation matrix is ​​used to determine how to effectively distribute the excitation signals to each actuator 1033. The matrix is ​​designed considering the structural characteristics of the aircraft 103, control requirements, and test objectives, ensuring that the signals received by each actuator 1033 reflect its role and function within the aircraft structure. In this way, the allocation matrix optimizes signal distribution, enabling the system to accurately test the aircraft's dynamic response in a simulated environment.

[0157] During signal distribution, the flight controller 1031, guided by the distribution matrix, distributes the generated excitation signals to each actuator 1033. Each actuator 1033 performs corresponding actions based on the received signals, including adjusting control surfaces, changing thrust, or performing other control tasks. Through these actions, the system can simulate the dynamic response of the aircraft under different operating conditions.

[0158] In one specific implementation, the allocation matrix consists of a two-dimensional array, where rows and columns correspond to excitation signals and actuators, respectively. Each matrix element represents a scaling factor used to determine the degree to which a specific excitation signal is allocated to a specific actuator. When designing the signal allocation matrix, the structural characteristics and control requirements of the aircraft must be considered, including the physical location and function of the actuators and their impact on the overall system response. By analyzing the system's feedback data, the signal allocation matrix can be dynamically adjusted to adapt to different test conditions and objectives. For example, when attitude adjustments in a specific direction are required, the matrix can be designed to allocate more excitation signals to the relevant actuators. This ability to carefully design and adjust not only improves the system's flexibility and accuracy but also provides strong support for optimizing the aircraft's control performance, ensuring optimal performance under various operating conditions.

[0159] The aircraft structure control coupling test system provided in this embodiment achieves intelligent distribution of excitation signals through the flight control system's functions. Specifically, the flight control system precisely distributes the excitation signals to each actuator according to a preset distribution matrix. This process ensures that each actuator receives appropriate signal strength and characteristics based on its role and function within the aircraft structure. In this way, the system can not only optimize the aircraft's control performance under various operating conditions but also improve the effectiveness and accuracy of the test, providing important reference data for aircraft design and optimization.

[0160] In one possible implementation, the flight controller 1031 in the aircraft structure control coupling test system 200 is also used to generate a stabilization control signal based on the excitation response and superimpose it onto the excitation signal during closed-loop testing.

[0161] Closed-loop testing is a real-time feedback control testing method designed to optimize system performance through continuous monitoring and adjustment. In this process, the aircraft structure control coupling test system 200 continuously monitors its own output parameters, such as attitude, velocity, and acceleration, and adjusts the input signals based on these parameters to ensure the aircraft achieves its expected performance goals.

[0162] Specifically, in the closed-loop test, the flight controller 1031 receives excitation responses from the control device 101. These excitation responses are the dynamic reactions of the aircraft 103 to the input excitation signals, reflecting the system behavior under the current flight conditions. The flight controller 1031 analyzes these excitation responses and uses a preset algorithm to identify the stability state and potential deviations of the aircraft 103. 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 algorithm, and it can be adjusted and determined according to actual needs.

[0163] Based on in-depth analysis of the excitation response, the Flight Control 1031 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 the stability augmentation control signal involves adjusting the control gain, changing control law parameters, or introducing additional control inputs to counteract instability factors.

[0164] 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.

[0165] The aircraft structure-control coupling test system provided in this embodiment enhances the system's stability and responsiveness by generating a stabilization control signal based on the excitation response during closed-loop testing and superimposing it onto the excitation signal. Specifically, the system can adjust the aircraft's control input in real time to cope with dynamically changing flight conditions, ensuring that the aircraft maintains its expected performance under various operating environments. This not only improves the accuracy and reliability of the test but also provides valuable data support for the design and optimization of the aircraft, ensuring its safe and efficient operation in practical applications.

[0166] In one possible implementation, Figure 2Based on the aircraft structure control coupling test system 200 shown, a second switch is set on the output path of the excitation signal. The second switch is located before the excitation signal and the stabilization control signal are superimposed. The second switch is used to control the access or cut-off of the excitation signal.

[0167] Specifically, the second switch is positioned on the output path of the excitation signal generated by the flight controller 1031, specifically before the excitation signal and the stability augmentation control signal are superimposed. The core function of the second switch is to control the flow of the excitation signal, that is, to determine whether the excitation signal can enter the subsequent signal processing stage. The control device 101 is responsible for managing the state of the second switch, determining whether the second switch is on or off based on the system's operating status and experimental requirements.

[0168] Under normal operating conditions, the second switch remains closed, meaning the excitation signal can be transmitted unimpeded to the next stage and superimposed on the stabilization control signal. However, under certain specific test conditions or emergency situations, such as when system instability is detected, special test procedures are required, or to prevent the excitation signal from adversely affecting the system, the control device 101 can quickly disconnect the second switch. By disconnecting the second switch, the system can immediately interrupt the transmission of the excitation signal, thus preventing the excitation signal from affecting system stability at inappropriate times.

[0169] The aircraft structure control coupling test system provided in this embodiment achieves flexible control over the flow of the excitation signal by setting a second switch on the output path of the excitation signal. This design allows the system to efficiently process the excitation signal and conduct structure control coupling tests under normal operating conditions, ensuring the smooth superposition of the excitation signal and the stability augmentation control signal. Simultaneously, under specific test conditions or in case of emergencies, the transmission of the excitation signal can be quickly cut off to prevent it from adversely affecting system stability, thereby providing additional protection for the safe operation of the aircraft.

[0170] In one possible implementation, Figure 2 Based on the aircraft structure control coupling test system 200 shown, a third switch is set on the output path of the stabilization control signal. The third switch is located before the excitation signal and the stabilization control signal are superimposed. The third switch is used to control the test state to be either open-loop or closed-loop.

[0171] Specifically, the third switch is placed on the output path of the stabilization control signal, before the excitation signal and the stabilization control signal are superimposed, so as to directly affect the transmission of the stabilization control signal and thus control the overall feedback mechanism of the system.

[0172] In open-loop mode, the third switch is open, and the stabilization control signal is not transmitted to the system's feedback path. In this state, the system's response depends only on the excitation signal and is unaffected by the stabilization control signal. This mode is typically used to test the system's natural response characteristics and analyze its behavior without feedback control.

[0173] In closed-loop mode, the third switch remains closed, allowing the stabilization control signal to be superimposed on the excitation signal and fed back into the system. In this state, the system response is regulated by the stabilization control signal, enabling automatic adjustments to maintain stability and optimize performance. Closed-loop mode is typically used to verify the system's control strategy and stability.

[0174] The third switch is managed by control device 101. Control device 101 determines the on / off state of the third switch based on whether the test requires an open-loop or closed-loop operation. This flexible control mechanism allows the system to switch states under different test conditions to meet various testing and analysis needs.

[0175] The aircraft structure control coupling test system provided in this embodiment achieves flexible switching between open-loop and closed-loop states by setting a third switch on the output path of the stabilization control signal. This design not only enhances the system's adaptability, enabling it to quickly adjust its state according to test requirements, but also improves the diversity and accuracy of testing.

[0176] In one possible implementation, there are multiple vibration acquisition devices 102, which are respectively located on the inertial navigation system 1032 of the aircraft 103, each actuator 1033, and at least one location characterizing the structural vibration mode.

[0177] Specifically, in the structural control coupling test of the aircraft 103, in order to accurately evaluate the vibration characteristics of the aircraft 103, it is necessary to select some representative parts as vibration monitoring points. Therefore, this application installs multiple vibration acquisition devices 102 at different locations on the aircraft 103 to achieve comprehensive vibration information acquisition, providing a data basis for determining whether the structural control coupling test has passed.

[0178] Among them, multiple vibration acquisition devices 102 are respectively installed on the inertial navigation system 1032 of the aircraft 103, each actuator 1033 and at least one position characterizing the vibration mode of the structure.

[0179] Specifically, the inertial navigation system 1032 is responsible for providing dynamic data such as linear acceleration and angular velocity of the aircraft 103. Vibration acquisition devices 102 are deployed at these locations to monitor the vibration response of various parts of the aircraft 103 in real time and perform correlation analysis with the vibration information provided by the inertial navigation system 1032 to ensure the structural stability of the aircraft 103 during motion.

[0180] The actuator 1033 is a component that generates corresponding actions based on the excitation signals from the flight controller 1031. A vibration acquisition device 102 is installed on each actuator 1033 to capture the vibration response of these excitation actions to the aircraft's airframe structure in real time.

[0181] Furthermore, considering that different parts of the aircraft 103 have different vibration characteristics, which affect the overall performance of the aircraft 103, a vibration acquisition device 102 is installed at at least one location of the aircraft 103 that characterizes the structural vibration mode, so that vibration information of these parts under different excitation signals can be collected.

[0182] Preferably, at least one location characterizing the structural vibration includes: the wingtip and / or tailtip of the aircraft 103.

[0183] Specifically, wingtips and tailtips are typical structural components of aircraft, typically experiencing complex vibration modes. The wings and tail are crucial aerodynamic control surfaces for aircraft 103, responsible for generating lift and maintaining stability during flight. When aircraft 103 is subjected to excitation, the wingtips and tailtips are susceptible to aerodynamic effects, especially during high-speed flight or aggressive maneuvers. The vibration characteristics of these components often significantly impact the overall structural stability of the aircraft.

[0184] Choosing the wingtip and / or tailtip as vibration information acquisition points allows for better capture of the vibration patterns of the aircraft 103 under different flight conditions.

[0185] During the structural control coupling test of the aircraft 103, the vibration acquisition equipment 102 installed at these key locations can collect vibration information of the aircraft wingtip and / or tailtip in real time, and promptly detect possible structural resonance problems or unfavorable vibration modes.

[0186] The aircraft structural control coupling test system provided in this embodiment aims to capture the main vibration modes of the aircraft under different flight conditions by monitoring vibrations at key locations such as the wingtips and / or tailtips. These locations, influenced by aerodynamic effects and flight dynamics, accurately reflect the structural vibration characteristics of the aircraft. By collecting vibration information from these key locations, the structural control coupling performance of the aircraft can be effectively evaluated, the design optimized, vibration resistance improved, and the stability and safety of the aircraft ensured during testing, thereby providing data support for the overall performance improvement of the aircraft.

[0187] By installing vibration acquisition devices 102 on the inertial navigation system 1032, each actuator 1033, and at least one location characterizing the structural vibration mode of the aircraft 103, the system can comprehensively collect vibration information of various key parts of the aircraft during the test, providing a data basis for determining whether the structural control coupling test is successful.

[0188] The aircraft structure-control coupling test system provided in this embodiment, by installing multiple vibration acquisition devices at the aircraft's inertial navigation system, actuators, and at least one location characterizing structural vibration modes, can comprehensively monitor the vibration characteristics of each component and accurately capture the coupling effect between the structure and the control system. This distributed vibration data acquisition method not only helps analyze the dynamic response of the aircraft under different operating conditions but also enhances the universality of the test, adapting to different types and configurations of aircraft and providing effective data support across platforms and models. Through the analysis of diverse vibration information, the system can provide data basis for aircraft optimization design, control system tuning, and fault early warning, enhancing the versatility and applicability of the test, thereby ensuring the safety, stability, and performance improvement of the aircraft.

[0189] Next, the specific location of the actuator 1033 will be determined according to the different types of aircraft 103.

[0190] Specifically, during the aircraft structure control coupling test, the location of actuator 1033 is determined according to the type of aircraft. Aircraft types include, but are not limited to, fixed-wing aircraft, rotorcraft, and compound-wing aircraft. Different aircraft types have different requirements and structures for their control systems, so the location of the actuator needs to be determined according to the aircraft type.

[0191] In one possible implementation, if the aircraft 103 is a fixed-wing aircraft, at least one actuation mechanism 1033 includes servos respectively located at the elevator, rudder, and aileron positions of the fixed-wing aircraft. Alternatively, in another case, at least one actuation mechanism 1033 includes servos respectively located at the elevator, rudder, and aileron positions of the fixed-wing aircraft.

[0192] Specifically, during the aircraft structure control coupling test, if the type of aircraft 103 is a fixed-wing aircraft, at least one actuator 1033 includes servo motors respectively located at the elevator, rudder and aileron positions of the fixed-wing aircraft.

[0193] For a V-tailed aircraft, the at least one actuator 1033 includes an elevator rudder and a servo motor respectively located on the V-tail and at the aileron positions.

[0194] Specifically, ailerons are used to control the roll of an aircraft, that is, its rotation around its longitudinal axis. Ailerons are usually mounted on the wings and can enable the aircraft to roll or maintain balance during flight.

[0195] The elevator is a control surface on the V-tail of a V-tailed aircraft, combining the functions of an elevator and a rudder. It is mainly used to control the pitch and yaw of the aircraft.

[0196] In fixed-wing aircraft, actuators are located at the positions of critical control surfaces such as elevators, rudders, and ailerons. In V-tail aircraft, actuators are located at critical control surfaces such as elevators, rudders, and ailerons. This is to ensure that the aircraft can accurately control its pitch, yaw, and roll motions. Each control surface is driven by an actuator and adjusts the aircraft's attitude and heading according to excitation signals.

[0197] 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.

[0198] This embodiment of the aircraft structure control coupling test system mainly describes the placement of actuators when the aircraft type is a fixed-wing aircraft. In this case, at least one actuator is respectively located at the elevator, rudder, and aileron positions of the fixed-wing aircraft. This configuration enables independent and coordinated control of each control surface, thereby accurately simulating the dynamic response of the aircraft during the test. Simultaneously, this design can effectively test the performance of the aircraft under the coupling effects of different control surfaces, thus helping to optimize the aircraft's structure and control system.

[0199] In one possible implementation, if the aircraft 103 is a rotorcraft, at least one actuator 1033 includes a motor that drives the propeller.

[0200] Specifically, during the aircraft structure control coupling test, if the type of aircraft 103 is a rotorcraft, at least one actuator 1033 includes a motor that drives the propeller.

[0201] Rotorcraft typically rely on multiple rotating propellers to provide lift and thrust, thereby achieving flight control. Therefore, the actuators of rotorcraft differ from those of fixed-wing aircraft, primarily concentrated on the motors that drive the propellers. These motors control the propeller speed and direction of rotation, thus controlling the aircraft's motion.

[0202] It is worth noting that the function of the motor driving the propeller is to adjust the propeller's speed and angle according to the instructions of the flight controller 1031, thereby controlling the aircraft's attitude and motion. Specifically, by precisely adjusting the motor's output power, the propeller's thrust can be affected, thereby controlling the aircraft's climb, roll, pitch, and yaw.

[0203] If the aircraft 103 is a rotorcraft, at least one actuator 1033 is directly connected to the propeller. Its function is to precisely adjust the propeller's rotational speed and angle to achieve precise control of the aircraft. This design provides high control accuracy and response speed, ensuring that the aircraft can quickly adapt to various dynamic changes during flight. If the actuator is designed far from the propeller, it may increase signal transmission delay and increase the weight and complexity of the system, affecting the aircraft's response performance. Therefore, at least one actuator 1033 includes a motor that drives the propeller, which helps reduce transmission delay and improve control immediacy, enabling the aircraft to quickly and accurately adjust its attitude in various flight conditions.

[0204] This embodiment provides an aircraft structure control coupling test system, primarily describing the placement of actuators when the aircraft type is a rotorcraft. In this case, at least one actuator is respectively located at the position of each propeller drive motor of the rotorcraft. This configuration allows for independent and coordinated control of the rotational speeds of each propeller, thereby accurately simulating the dynamic response of the aircraft during testing. This design can effectively test the performance of the aircraft under different propeller thrust and speed adjustments, especially the coupling effect during the aircraft's climb, pitch, roll, and yaw movements.

[0205] In one possible implementation, if the type of aircraft 103 is a compound wing aircraft, at least one actuator 1033 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.

[0206] Specifically, during the aircraft structure control coupling test, if the type of aircraft 103 is a compound wing aircraft, at least one actuator 1033 includes: a servo motor located at the elevator, rudder or at the elevator rudder position, a servo motor at the aileron position, and a motor driving the propeller.

[0207] Among them, compound wing aircraft are aircraft that use composite material structures. Their wing design is not limited to the traditional single wing shape, but integrates a variety of different types or functions of wing structure to improve flight performance, reduce weight and enhance the stability and maneuverability of the aircraft.

[0208] Considering the multi-wing configuration and varying lift distribution of the compound wing aircraft, at least one actuator 1033 includes: a servo located at the elevator, rudder, or elevator / rudder position, and a servo located at the aileron position. The servos at the elevator, rudder, and aileron positions are used to control the pitch, yaw, and roll motions of the aircraft, respectively.

[0209] Specifically, the elevator controls the pitch angle of the aircraft to ensure that it can fly up and down correctly.

[0210] The rudder controls the yaw angle, enabling the aircraft to maintain a stable course.

[0211] The ailerons control the aircraft's roll motion, ensuring that it can turn smoothly or maintain its flight attitude.

[0212] The elevator is a control surface on the V-tail of a V-tailed aircraft, combining the functions of an elevator and a rudder. It is mainly used to control the pitch and yaw of the aircraft.

[0213] In the aircraft structure control coupling test, at least one actuator 1033 includes a servo motor located at the elevator, rudder, or elevator rudder position. The servo motor at the aileron position can precisely control the attitude of the aircraft to ensure the stability and maneuverability of the aircraft in flight.

[0214] The elevators, rudders, and ailerons on a compound wing aircraft work together to effectively adjust the aircraft's pitch, yaw, and roll, thereby improving the aircraft's maneuverability and stability.

[0215] The elevators and ailerons on a V-tail aircraft work in tandem to effectively regulate pitch, yaw, and roll. Due to the unique structure of the V-tail, the elevator typically combines the functions of an elevator and a rudder. By adjusting the servo's movement, not only can the pitch angle be controlled, but yaw can also be adjusted simultaneously. The ailerons control the roll. Through precise coordination of the servos, V-tail aircraft can achieve accurate control of attitude and heading with a simplified control system, thereby improving the aircraft's maneuverability, stability, and flight efficiency.

[0216] Meanwhile, compound-wing aircraft are typically equipped with an electric motor-driven propeller as their propulsion system. Especially in vertical takeoff and landing (VTOL) flight mode, the motor needs to adjust thrust to achieve vertical takeoff and landing. In horizontal flight mode, the motor adjusts thrust to maintain a stable flight speed. There is a coupling relationship between the motor and the aircraft structure; changes in motor thrust can cause changes in the stress on the aircraft structure, thus affecting the aircraft's dynamic response. Therefore, in the aircraft structure control coupling test, at least one actuator 1033 also includes a motor driving the propeller, which can precisely adjust thrust to achieve vertical takeoff and landing and horizontal flight, maintaining the stability and high efficiency of the aircraft in different flight modes.

[0217] This embodiment of the aircraft structure control coupling test system mainly describes the placement of actuators when the aircraft type is a compound wing aircraft. In this case, at least one actuator is respectively located at each control surface and propeller drive motor position of the aircraft. With this configuration, the system can precisely control each control surface of the aircraft to adjust the aircraft's attitude and stability, while independently adjusting the rotational speed and thrust of each propeller, simulating the impact of thrust changes on the aircraft's dynamic response under different flight conditions. This helps to test the performance of the compound wing aircraft under different flight modes, especially the coupling effect between various control systems in complex aerodynamic environments, thereby ensuring that the aircraft has strong stability and maneuverability under various flight conditions.

[0218] In one possible implementation, Figure 2 Based on the aircraft structure control coupling test system 200 shown, the aircraft structure control coupling test system 200 also includes a support device.

[0219] The main function of the support device is to support the aircraft during the test, enabling it to be tested stably in different configurations.

[0220] Specifically, the support device can adjust its support method according to the different attitudes or operating states of the aircraft to ensure that the aircraft can maintain the preset configuration under various experimental conditions. Through this design, the aircraft can conduct control coupling experiments in different flight modes, ensuring the ability to study the dynamic response of the aircraft under various configurations, the effectiveness and stability of the control system, and the impact of structural deformation on control performance. At the same time, the addition of the support device makes the aircraft more stable during testing, avoiding data errors or test anomalies that may be caused by unstable external support, thereby improving the reliability and accuracy of the test results.

[0221] In one specific embodiment, 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 for open-loop and closed-loop tests of takeoff and landing configurations.

[0222] Specifically, the elastic suspension system simulates the vibration response of an aircraft under different flight configurations by suspending it. This system is primarily used for open-loop and closed-loop tests of aircraft under various configurations, including takeoff, landing, transition, and cruise. Utilizing the elastic suspension system, the dynamic response of the aircraft can be realistically reproduced, especially by avoiding interference from ground contact during testing, thereby accurately measuring the stability, control performance, and vibration characteristics of the aircraft under various flight configurations.

[0223] Takeoff and landing configuration refers to the configuration in which an aircraft climbs by vector power from a stationary ground or descends by vector power during landing. The main actuators of the aircraft are direct force actuators (generally rotors, ducted fans, etc.).

[0224] Transition configurations refer to the transition modes between rotor and fixed-wing modes. These include tilt-rotor transition configurations and dual-rotor transition configurations.

[0225] 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.).

[0226] It should be noted that aircraft configuration also includes aircraft equipment status configuration. Because the aircraft equipment status configuration covers 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.

[0227] It is worth noting that open-loop and closed-loop tests must be conducted separately for different flight configurations, with open-loop tests preceding closed-loop tests. Open-loop tests are performed first because they can independently test the coupling characteristics of the aircraft structure and control system without feedback control interference, providing basic data for subsequent closed-loop control debugging. Closed-loop tests, on the other hand, can verify the dynamic response and adaptability of the aircraft control system under different configurations after feedback is introduced.

[0228] Before conducting the test, the support device should be selected appropriately based on the aircraft's weight and lowest-order resonant frequency. Optional support devices for the aircraft include elastic suspension systems and support frames.

[0229] Specifically, the elastic suspension device is an air spring or an elastic rope.

[0230] 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.

[0231] 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 evaluation under various test conditions. In one specific implementation, the elastic rope is a rubber rope, and the support equipment for rubber rope suspension mainly consists of the rubber rope 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.

[0232] 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.

[0233] The support frame functions differently from the elastic suspension device. It is mainly used to support the aircraft in open-loop and closed-loop tests of take-off and landing configurations, ensuring that the aircraft can carry out various structural control coupling tests in a confined environment.

[0234] 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.

[0235] 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.

[0236] The tests should include tests under various load conditions, including but not limited to open-loop and closed-loop tests.

[0237] Specifically, open-loop testing, also known as open-loop stability margin reserve testing, typically uses a swept-frequency signal or a single-point signal to excite each channel and performs 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. In open-loop testing, the flight control system does not make feedback adjustments, testing its dynamic response without external intervention, and testing the preliminary coupling characteristics of the aircraft structure and control system. At this time, the relationship between the aircraft's control input and its response is determined by factors such as the stiffness of the structure itself, mass distribution, and aerodynamic effects. The support frame ensures that the aircraft maintains the correct attitude and positioning during the test, avoiding external disturbances from affecting the accuracy of the data.

[0238] Closed-loop testing, also known as closed-loop stability margin verification, typically involves pulse or sweep frequency excitation of each channel followed by time-domain analysis. During the test, the sweep frequency amplitude is gradually changed to verify the stability margin reserves of each channel. Closed-loop testing is conducted after incorporating feedback adjustment from the flight control system. The flight control system adjusts control commands based on real-time data, thereby optimizing flight attitude and stability in real time during takeoff and landing. Through this test, personnel can observe how the aircraft responds to external disturbances during takeoff and landing, assess the robustness and adaptability of the control system, and verify the coupling effect between the structure and the control system. The support frame not only provides necessary fixed support in this process but also simulates the stress conditions of the aircraft under different loads and external environments during testing, ensuring the accuracy and reliability of the test data. This provides crucial information for optimizing the aircraft's takeoff and landing performance and adjusting control strategies.

[0239] A pulse signal is a rapidly changing signal over a short period of time. Its characteristics include a very short duration and a large amplitude variation within a certain timeframe, which can be used to stimulate the frequency response of an aircraft. A swept-frequency signal, on the other hand, covers multiple frequency ranges to test the aircraft's response characteristics at different frequencies.

[0240] 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.

[0241] 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.

[0242] 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.

[0243] The selection of specific test methods should be based on the progress of research and development and the test requirements.

[0244] 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.

[0245] 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.

[0246] Optionally, the support frame can be a specially designed support frame or the landing gear of an aircraft.

[0247] 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.

[0248] 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.

[0249] 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.

[0250] 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.

[0251] The aircraft structure-control coupling test system provided in this embodiment, through the combination of an elastic suspension device and a support frame, can effectively support open-loop and closed-loop tests under different flight configurations. The elastic suspension device suspends the aircraft, facilitating open-loop and closed-loop tests in takeoff and landing, transition, and cruise configurations. This helps to accurately simulate the dynamic behavior and response characteristics of the aircraft under different flight states. The support frame is specifically used to support the aircraft in takeoff and landing configuration tests, ensuring the aircraft remains stable and fixed during testing, avoiding the influence of external disturbances on the data. By implementing this system, the structural and control coupling characteristics of the aircraft under different configurations can be comprehensively tested, optimizing the aircraft's control strategy and design, and improving the aircraft's takeoff and landing performance and stability.

[0252] In one possible implementation, the frequency of the support device is lower than a preset frequency threshold for the first-order structural mode frequency of the aircraft.

[0253] Specifically, the first-order structural modal frequency of an aircraft refers to the first natural vibration frequency of the aircraft when it is subjected to external forces 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 that it is not subject to any external constraints.

[0254] 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.

[0255] 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.

[0256] 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.

[0257] In one specific implementation, the frequency domain threshold is one-third of the first-order structural mode frequency.

[0258] Specifically, by setting the frequency of the support device to one-third of the first-order structural mode frequency of the aircraft, vibration resonance can be effectively avoided. This design not only ensures the stability of the aircraft under various test conditions but also provides sufficient safety margins to ensure the accuracy and reliability of test data. Simultaneously, this frequency ratio avoids vibration interference without affecting the working efficiency of the support device, contributing to obtaining accurate test data, optimizing the aircraft's design and control strategies, and improving the aircraft's performance and stability.

[0259] In one possible implementation, Figure 2 Based on the aircraft structure control coupling test system 200 shown, the aircraft structure control coupling test system 200 also includes a control surface elastic fixture and / or counterweight, which are used to increase the excitation force of the control surface during frequency sweep; the control surface elastic fixture is set on the control surface of the aircraft; the counterweight is set on the rear rafter of the control surface of the aircraft.

[0260] 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.

[0261] The use of elastic fixtures on control surfaces can increase the elastic modulus and response characteristics of the control surfaces, resulting in a more significant vibration response at certain frequencies. This is particularly important for frequency sweep tests, which typically cover a wide frequency range and use excitations at different frequencies to study the coupling effect between the aircraft structure and control system. The role of the elastic fixtures is to enhance the stiffness and deformation capacity of the control surfaces, thereby increasing the response amplitude of the control surfaces to excitations at different frequencies during frequency sweeps. This helps test personnel to more clearly observe and analyze the dynamic behavior of the control surfaces, especially under low-frequency and high-frequency excitations, enabling more precise capture of the interaction between the control system and the structure.

[0262] The use of counterweights effectively simulates the mass distribution and aerodynamic disturbances that aircraft may encounter during actual flight, ensuring that experimental data more closely reflects reality and improving the reliability and applicability of the tests. Specifically, the main purpose of counterweights is to further regulate the interaction between the control surface and the airflow by altering the mass distribution of the control surface. The counterweights are placed at the rear sill of the control surface because changes in the mass of the rear sill have a significant impact on the aerodynamic forces of the control surface. The addition of counterweights effectively changes the position of the control surface's center of mass, thereby altering its vibration characteristics and resulting in a stronger vibration response during frequency sweeps. Especially near some critical frequencies, the control surface experiences greater excitation forces, which helps improve the sensitivity and accuracy of the tests. In this way, counterweights can help simulate the aerodynamic disturbances that control surfaces may encounter during actual flight, thus providing a more realistic dynamic response.

[0263] By combining elastic tooling on the control surfaces with counterweights, the vibration response of the control surfaces can be improved during frequency sweep tests, making the dynamic characteristics of the aircraft more prominent at different frequencies. This helps to capture the coupling effect between the aircraft's structure and control system under a wide range of frequency excitations, especially in the alternating changes between low-frequency and high-frequency vibrations of the aircraft, where the transmission and interaction of vibrations can be clearly observed.

[0264] Meanwhile, the elastic tooling and counterweights on the control surfaces, by increasing the excitation force, help test personnel evaluate the performance of the aircraft control system under various dynamic environments, especially under large amplitude or critical frequency conditions, and can better analyze the response behavior of the aircraft under complex vibrations.

[0265] Furthermore, the design of the control surface's elastic tooling and counterweights can also help avoid resonance. By precisely adjusting the vibration characteristics of the control surface, its natural frequency and the frequency of the external excitation source can be kept within a reasonable range, preventing resonance from causing potential damage to the aircraft structure.

[0266] The aircraft structure-control coupling test system provided in this embodiment, through the design of elastic tooling on the control surfaces and / or counterweights, can not only effectively enhance the excitation force in frequency sweep tests, but also more accurately simulate the dynamic behavior of the aircraft under various vibration environments. This design helps to provide more accurate test data on the coupling characteristics of the aircraft structure and control system, thereby providing important support for the optimized design and performance improvement of the aircraft.

[0267] In one possible implementation, the control device 101 is also used to control the flight controller 1031 to drive the configuration switching of the aircraft 103.

[0268] In structural control coupling tests of aircraft, the configuration switching of aircraft 103 has a crucial impact on the test results. The cooperation between control equipment 101 and flight control system 1031 is the core of achieving this switching, ensuring the stability and performance of the aircraft under different flight conditions. Through comprehensive testing of different configurations, structural control coupling tests can evaluate and optimize the responsiveness and adaptability of the control system.

[0269] Specifically, the control device 101 controls the flight controller 1031 through configuration switching commands to realize configuration switching of the aircraft 103 in various flight modes.

[0270] Aircraft are classified 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.

[0271] 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.).

[0272] Transition configurations refer to the transition modes between rotor and fixed-wing modes. These include tilt-rotor transition configurations and dual-rotor transition configurations.

[0273] 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.).

[0274] 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.

[0275] The aircraft structure-control coupling test system provided in this embodiment enables configuration switching of an aircraft across multiple flight modes by precisely coordinating control equipment and flight control. This design enhances the aircraft's maneuverability and response speed under different flight conditions, while ensuring the stability and safety of the switching process, avoiding potential instability or unexpected situations. Through testing various configurations, the system can deeply evaluate and optimize the adaptability and responsiveness of the control system, thereby enhancing the overall performance of the aircraft. Furthermore, this design has broad applicability, adapting to different types of aircraft and diverse mission requirements, providing important reference and data support for future flight control system development.

[0276] It is worth noting that for rotorcraft or compound wing aircraft, high-energy rotating propellers must be driven during structural control coupling tests to effectively excite the control channels and thus stimulate 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 rotational speed as in actual flight. 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 rotorcraft structural control coupling tests, a suitable propeller must be carefully selected as the excitation source to ensure the safety and effectiveness of the test.

[0277] In one possible implementation, the propeller blades of a rotorcraft or compound wing aircraft can be used as the excitation source, especially for aircraft with larger propellers, where collective pitch needs to be reduced. 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.

[0278] 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.

[0279] 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):

[0280]

[0281]

[0282] 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 differential speed of the propeller. 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.

[0283] In one possible implementation, for rotorcraft or compound aircraft, an open-loop test is conducted using the selected propeller as the excitation source. For fixed-wing aircraft, an open-loop test can be conducted directly. Based on the results of the open-loop structure control coupling test for 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 based on the open-loop test data. Formula (3) is shown below:

[0284]

[0285] 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.

[0286] The aircraft structure control coupling test system provided in this embodiment, targeting rotorcraft and compound wing aircraft, precisely selects the size and characteristics of the propeller using a specific formula, thereby ensuring that the control excitation state is consistent with actual flight conditions during the test. Specifically, the precise propeller selection not only improves the effectiveness of the test but also ensures the safety and stability of the excitation source at high speeds. Furthermore, the selected propeller can effectively excite the modal characteristics of the aircraft structure, providing accurate dynamic response data, thus helping to optimize flight control design. By reducing the risks of high-speed operation, this design further enhances the safety and operability of the test.

[0287] The following section uses an aircraft as an example to elaborate on the aircraft structure control coupling test system proposed in this application.

[0288] Figure 3This is a schematic diagram of Embodiment 2 of the aircraft structure control coupling test system provided in this application. Figure 3 As shown, the aircraft structure control coupling test system includes a main control PC 301, an aircraft structure 300, a flight control system 1031, an inertial navigation system 1032, an actuator 1033, and a vibration acquisition device 102. It is worth noting that the main control PC 301 is the aforementioned control device 101.

[0289] The aircraft structure 300 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.

[0290] The aircraft structure 300, as the controlled object, has actuators installed on its fuselage structure, namely, inertial navigation system 1032, actuator 1033, and vibration acquisition device 102. It is worth noting that if the aircraft type is a fixed-wing aircraft, the actuators for this type of aircraft are mainly servo motors installed in the elevator, rudder, and aileron positions. If the aircraft type is a rotary-wing aircraft, the actuators for this type of aircraft are mainly motors installed on the arms to drive the propellers. If the aircraft type is a compound-wing aircraft, the actuators for this type of aircraft are mainly servo motors installed in the elevator, rudder, and aileron positions, as well as motors to drive the propellers.

[0291] During aircraft operation, the pilot / flight controller uses loop-controlled actuators such as control surfaces and motors (1033) 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.

[0292] During the test, the aircraft structure 300 requires a special elastic support device for support, 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.

[0293] The actuator 1033 communicates directly with the flight controller 1031 via a serial port / network port and receives control command signals from the flight controller 1031.

[0294] The inertial navigation system 1032 also communicates directly with the flight control system 1031 via a serial port, feeding back the measured aircraft status and vibration information to the flight control system 1031.

[0295] Vibration acquisition equipment 102 is installed at different locations on the aircraft fuselage. Optionally, the specific location of the vibration acquisition equipment 102 on the aircraft fuselage can be determined according to the requirements of the ground vibration test. It is worth noting that vibration acquisition equipment 102 must be installed at the mounting locations of inertial navigation system 1032 and motors to allow test personnel to collect and analyze vibration information of the aircraft structure.

[0296] The flight control computer 1031 is connected to the actuator 1033, the inertial navigation system 1032, and the main control PC 301.

[0297] The Flight Control 1031 computer consists of two parts: computer hardware and software. The software includes flight control software and airborne software. The flight control software provides stability augmentation control law functions, 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 each actuator, specifically including fixed-wing control surface commands, rotorcraft motor throttle commands, and helicopter pitch control servo commands. The airborne software provides frequency sweep signal generation functions, enabling the generation of experimental frequency sweep signals.

[0298] The main control PC301 is connected to the flight control 1031 computer via a network port. The main control PC301 has host computer software installed, which allows configuration of the signal waveform information generated by the onboard 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 onboard software of the flight control 1031, and can also collect and monitor aircraft vibration information acquired by the inertial navigation system 1032.

[0299] In addition, the main control PC301 is also connected to the vibration acquisition device 102, which can directly monitor and analyze the vibration information collected by the vibration acquisition device 102. At the same time, the main control PC301 also has attitude limit measures to ensure that it can automatically shut down after the vibration exceeds the limit.

[0300] 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.

[0301] Figure 4 This is a schematic diagram of Embodiment 3 of the aircraft structure control coupling test system provided in this application. Figure 4 As shown, the aircraft structure control coupling test system involves switch 1, switch 2, switch 3, measurement 1, measurement 2, measurement 3, allocation matrix, actuator, controlled object, sensor and stabilization control law.

[0302] Switch 1 (the first switch) is located between the flight controller and at least one actuator, specifically at the position after 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 damage to the aircraft structure.

[0303] Switch 2 (second switch) is set 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, and is used to control the access or cut-out of the sweep frequency control signal (excitation signal).

[0304] Switch 3 (the third switch) 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.

[0305] 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.

[0306] 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.

[0307] Measurement 3 is located after the sensor and is used to observe and record the sensor measurements.

[0308] based on Figure 3 and Figure 4 The aircraft structure control coupling test system shown can set the opening and closing of switch 3 via the main control PC301 host computer, thereby realizing 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 airborne software configured on the main control PC301 can generate information such as the frequency and amplitude of the excitation signal. The main control PC301 can also detect whether the excitation signal output by the flight controller 1031 meets the requirements from measurement point 1. After confirming that the signal is correct, the connection switch between the flight controller 1031 and the actuator 1033 is turned on.

[0309] 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 PC301, 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.

[0310] 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.

[0311] Figure 5 This is a schematic diagram of Embodiment 4 of the aircraft structure control coupling test system provided in this application. Figure 5 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 5 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.

[0312] Figure 6 This is a schematic diagram of Embodiment 5 of the aircraft structure control coupling test system provided in this application. Figure 6 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.

[0313] 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.

[0314] 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.

[0315] 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.

[0316] 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.

[0317] 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.

[0318] 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.

[0319] 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.

[0320] The test process is carried out according to the relevant provisions in the preset 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 conditions, completing the installation of 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 completing vibration signal acquisition by the inertial navigation system and vibration signal acquisition by the 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.

[0321] 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.

[0322] 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.

[0323] 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 coupling test system for aircraft structural control, characterized in that, include: Control equipment, and vibration acquisition equipment installed on the aircraft's airframe structure; The control device is connected to the flight controller of the aircraft and is used to configure the frequency range corresponding to the test in the flight controller. The flight controller is used to generate an excitation signal according to the frequency range and output it to at least one actuator. The inertial navigation system and the vibration acquisition device of the aircraft are used to collect vibration information on the airframe structure of the aircraft during execution; The control device is also used to determine whether the structural control coupling test has passed based on the vibration information.

2. The system according to claim 1, characterized in that, The control device is also used to configure at least one of the following: the generation, termination, cancellation, frequency, and amplitude of the excitation signal.

3. The system according to claim 1, characterized in that, The control device is specifically used to determine the excitation response based on the vibration information, and to determine whether the structural control coupling test passes based on the excitation response.

4. The system according to claim 3, characterized in that, The control device is specifically used to: determine that the structural control coupling test has passed when the excitation response is greater than the preset amplitude-frequency margin index, and determine that the structural control coupling test has failed when the excitation response is less than or equal to the amplitude-frequency margin index.

5. The system according to claim 4, characterized in that, The amplitude-frequency margin index is 6dB or 9dB.

6. The system according to any one of claims 1 to 5, characterized in that, A first switch is provided between the flight controller and the at least one actuator; The control device is also used to control the on / off state of the first switch.

7. The system according to any one of claims 1 to 5, characterized in that, The flight control system is also used to distribute the excitation signal to each actuator according to a preset distribution matrix.

8. The system according to any one of claims 1 to 5, characterized in that, The flight control system is also used to generate a stabilization control signal based on the excitation response and superimpose it onto the excitation signal during closed-loop testing.

9. The system according to claim 8, characterized in that, A second switch is provided on the output path of the excitation signal. The second switch is located before the excitation signal and the stabilization control signal are superimposed. The second switch is used to control the access or cut-off of the excitation signal.

10. The system according to claim 8, characterized in that, A third switch is provided on the output path of the stabilization control signal. The third switch is located before the excitation signal and the stabilization control signal are superimposed. The third switch is used to control the state of the test to be either open-loop or closed-loop.

11. The system according to any one of claims 1 to 5, characterized in that, The number of vibration acquisition devices is multiple; Multiple vibration acquisition devices are respectively installed at the inertial navigation system, each actuator, and at least one location representing the vibration mode of the structure.

12. The system according to claim 11, characterized in that, The location of at least one characterizing structural vibration includes: the wingtip and / or tailtip of the aircraft.

13. The system according to any one of claims 1 to 5, characterized in that, The aircraft is a fixed-wing aircraft, and the at least one actuator includes: servo motors respectively disposed at the elevator, rudder and aileron positions of the fixed-wing aircraft, or servo motors respectively disposed at the elevator rudder and aileron positions of the fixed-wing aircraft.

14. The system according to any one of claims 1 to 5, characterized in that, The aircraft is a rotorcraft, and the at least one actuator includes a motor that drives the propeller.

15. The system according to any one of claims 1 to 5, characterized in that, The aircraft is a compound wing aircraft, and the at least one actuator includes: Servo mechanisms located at the elevator, rudder, and aileron positions, and motors driving the propeller; or, The servo motors are located at the positions of the elevator and ailerons, as well as the motors that drive the propeller.

16. The system according to any one of claims 1 to 5, characterized in that, The system also includes a support device for supporting the aircraft in different configurations during the test.

17. The system according to claim 16, characterized in that, The support device includes a flexible suspension device or a support frame; The elastic suspension device 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 in open-loop and closed-loop tests of takeoff and landing configurations.

18. The system according to claim 16, characterized in that, The frequency of the support device is lower than the preset frequency threshold of the first-order structural mode frequency of the aircraft.

19. The system according to claim 18, characterized in that, The frequency domain threshold is one-third of the frequency of the first-order structural mode.

20. The system according to any one of claims 1 to 5, characterized in that, The system also includes: a rudder surface elastic tooling and / or counterweight, used to increase the excitation force of the rudder surface during frequency sweep; The control surface elastic tooling is mounted on the control surface of the aircraft; The counterweight is located on the rear axle of the aircraft's control surface.

21. The system according to any one of claims 1 to 5, characterized in that, The control device is also used to control the flight controller to drive the configuration switching of the aircraft.

22. The system according to any one of claims 1 to 5, characterized in that, When the aircraft is a rotorcraft or a compound wing aircraft The propeller of the rotorcraft or compound wing aircraft is selected according to the following formula (1) or according to the following formulas (1) and (2): 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.

23. The system according to claim 4, characterized in that, The control device is specifically used for: 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.

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