A wind tunnel flutter test device

Abstract

A wind tunnel flutter testing device, relating to the field of wind tunnel testing technology, includes an airfoil model and a control unit. The airfoil model is fixedly connected to the wind tunnel interface by bolts. Linear displacement sensors and torque strain gauges are installed on the main beam of the airfoil model. Flight control sensors are installed on the root body of the airfoil model. A pulsating pressure sensor is installed on the lower skin surface of the airfoil model. An angle sensor is installed on the control surface shaft of the airfoil model. An acceleration sensor is installed on the stiffener of the airfoil model. The signal output pins of the linear displacement sensors, torque strain gauges, flight control sensors, pulsating pressure sensors, and angle sensors are all electrically connected to the input pins of the control unit. The output pins of the control unit are electrically connected to the servo motors of the airfoil model and a host computer, respectively. This device addresses the problem of single signal acquisition in traditional testing equipment.

Description

Technical Field

[0001] The utility model relates to the field of wind tunnel experimental technology, specifically to a wind tunnel flutter experimental device. Background Technology

[0002] Flutter is a self-excited vibration of a structure moving in airflow, caused by the interaction of aerodynamic, inertial, and elastic forces. Flutter in aircraft is particularly dangerous. When the flight speed exceeds the flutter critical speed, the amplitude and dynamic stress in the structure can increase dramatically, leading to rapid (sometimes within seconds) structural failure during flight. Therefore, any form of flutter is unacceptable. Accurately obtaining key flutter characteristic parameters such as the aircraft's flutter critical speed and flutter frequency is essential, and conducting flutter tests in a wind tunnel is currently the most reliable experimental method.

[0003] Existing wind tunnel flutter test equipment uses airfoil models with a fixed center of mass, and the aerodynamic characteristics of the airfoil surface are measured by several miniature pressure sensors arranged on the airfoil surface. However, this model is limited to acquiring only pressure signals. Although flutter results can be obtained in this way, the host computer also uses only a single physical algorithm model when analyzing the flutter generation. Furthermore, due to the noise in the experimental data, there is a certain difference between the ideal physical algorithm model and the actual experiment, resulting in low accuracy of the experimental results.

[0004] Chinese patent CN116754172A discloses a high Mach number free-flow wind tunnel flutter test system and test method, which uses accelerometers and strain gauges to detect flutter data of the model, but also has certain limitations.

[0005] Therefore, we propose a device that can acquire experimental signals from multiple angles. Utility Model Content

[0006] The purpose of this invention is to provide a wind tunnel flutter experimental device that solves the problem of single signal acquisition by traditional experimental devices.

[0007] This utility model is achieved through the following technical solution: A wind tunnel flutter testing device includes a wing segment model and a control unit. The wing segment model is fixedly connected to the wind tunnel interface by bolts. A linear displacement sensor and a torque strain gauge are installed on the main beam of the wing segment model. A flight control sensor is installed on the root body of the wing segment model. A pulsating pressure sensor is installed on the lower skin surface of the wing segment model. An angle sensor is installed on the control surface shaft of the wing segment model. An acceleration sensor is installed on the stiffener of the wing segment model. The signal output pins of the linear displacement sensor, torque strain gauge, flight control sensor, pulsating pressure sensor, and angle sensor are all electrically connected to the input pins of the control unit. The output pins of the control unit are electrically connected to the servo motor of the wing segment model and the host computer, respectively.

[0008] Furthermore, the control unit includes a housing, and an air switch, a filter, a switching power supply, a PLC, a servo driver, and an edge computing gateway installed inside the housing. The input terminal of the air switch is electrically connected to an external power supply, and the output terminal of the air switch is electrically connected to the switching power supply. The power output terminal of the switching power supply is electrically connected to the PLC and the edge computing gateway, and the input terminal of the PLC is electrically connected to the filter. The input terminal of the filter is electrically connected to a linear displacement sensor, a torque strain gauge, a flight control sensor, a pulsating pressure sensor, and an angle sensor, respectively. The output terminal of the PLC is electrically connected to the servo driver and the edge computing gateway, respectively. The servo driver is electrically connected to the servo motor of the wing segment model, and the edge computing gateway is electrically connected to the host computer.

[0009] Furthermore, the linear displacement sensor is model MTI-2100.

[0010] Furthermore, the number of torque strain gauges is two.

[0011] Furthermore, the model number of the torque strain gauge is CEA-06-250UW-350.

[0012] Furthermore, the flight control sensor is model ADIS16470.

[0013] Furthermore, the model number of the pulsating pressure sensor is XCQ-062.

[0014] Furthermore, the angle sensor is model RLS RM08.

[0015] The technical solution of this utility model has at least the following advantages and beneficial effects: This utility model discloses a wind tunnel flutter test device, which can detect various data when the airfoil model flutters through multiple sensors, that is, reflect the flutter test results of the airfoil model from multiple dimensions, and ensure the accuracy of the flutter test results.

[0016] In addition, the control unit can collect signals from multiple sensors, summarize them, and send them to the host computer for centralized data processing, thereby improving data processing speed. Attached Figure Description

[0017] Figure 1 This is a schematic diagram of the structure of this utility model; Figure 2 This is a schematic diagram of the installation structure of a control unit according to the present invention; Figure 3 This is a schematic diagram of a control unit connection structure according to the present invention.

[0018] Reference numerals: 1. Wing segment model; 101. Main beam; 102. Linear displacement sensor; 103. Torque strain gauge; 104. Root body; 105. Flight control sensor; 106. Lower skin; 107. Pulsating pressure sensor; 108. Control surface; 109. Angle sensor; 110. Servo motor; 111. Rib; 112. Acceleration sensor; 2. Control unit; 201. Housing; 202. Air switch; 203. Filter; 204. Switching power supply; 205. PLC; 206. Servo driver; 207. Edge computing gateway. Detailed Implementation

[0019] To make the objectives, technical solutions, and advantages of the embodiments of this utility model clearer, the technical solutions of the embodiments of this utility model will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this utility model, and not all embodiments. The components of the embodiments of this utility model described and shown in the accompanying drawings can generally be arranged and designed in various different configurations.

[0020] Example 1 like Figure 1-3The wind tunnel flutter test apparatus shown includes a wing segment model 1 and a control unit 2. The wing segment model 1 is fixedly connected to the wind tunnel interface by bolts. This wing segment model 1 is a scaled-down model 1 of an aircraft wing. After the wing segment model 1 is fixedly connected to the wind tunnel interface, wind tunnel blowing air onto the wing segment model 1 can begin the wind tunnel test. The main beam 101 of the wing segment model 1 is equipped with a linear displacement sensor 102 and a torque strain gauge 103. The linear displacement sensor 102 is used to detect the lateral dynamic displacement at the end of the main beam 101. Specifically, the model number of the linear displacement sensor 102 is... The MTI-2100 is used, while the torque strain gauge 103 is used to detect the torsional deformation stress of the main beam 101. In addition, the model of the torque strain gauge 103 is CEA-06-250UW-350. There are two torque strain gauges 103. Both torque strain gauges 103 can detect the torsional deformation stress of the main beam 101. However, when the host computer processes the signal, it analyzes the torque on the main beam 101 based on the output signal of one torque strain gauge 103, while the host computer analyzes the bending degree of the main beam 101 based on the output signal of the other torque strain gauge 103. The root body 104 of the wing segment model 1 is equipped with a flight control sensor 105, which is used to monitor the rigid body motion of the wing segment model 1 as a whole, and can distinguish between real flutter and interference signals from the vibration of the supporting structure, as well as provide an attitude reference to ensure the coordinate system consistency of other sensor data; in particular, the flight control sensor 105 is model ADIS16470. A pulsating pressure sensor 107 is installed on the surface of the lower skin 106 of the wing segment model 1. This sensor can detect high-frequency aerodynamic pressure fluctuations on the skin surface, thereby capturing unsteady aerodynamic phenomena such as airflow separation and shock wave oscillations. This data is used by the host computer to establish the phase relationship between pressure distribution and structural vibration, and then to identify the critical angle of attack / Mach number combination that leads to flutter. Specifically, the pulsating pressure sensor 107 is model XCQ-06. An angle sensor 109 is installed on the pivot of the control surface 108 of the wing segment model 1. This sensor can detect the real-time deflection angle of the control surface 108. After this data is transmitted to the PLC 205 of the subsequent control unit 2, the PLC 205 determines the difference between the real-time deflection angle of the control surface 108 and the deflection angle of the control surface 108 required by the experimental parameters, and controls the control surface 108 to deflect to the required angle value. Specifically, the angle sensor 109 is model RLS RM08. An acceleration sensor 112 is installed on the stiffener 111 of the wing segment model 1. This sensor is model Endevco. The 7264G-2000 accelerometer sensor 112 directly measures the vibration acceleration time-domain signal at the connection between the main beam 101 and the stiffening plate 111, thereby providing the real vibration boundary conditions of the physical model when the software of the host computer processes it in the subsequent process.

[0021] It should be noted that the experimental parameters are data that the staff have preset in advance.

[0022] The signal output pins of the linear displacement sensor 102, torque strain gauge 103, flight control sensor 105, pulsating pressure sensor 107, angle sensor 109, and acceleration sensor 112 are all electrically connected to the input pins of the control unit 2. The output pins of the control unit 2 are electrically connected to the servo motor 110 of the wing segment model 1 and the host computer, respectively. The control unit 2 transmits the output data of the linear displacement sensor 102, torque strain gauge 103, flight control sensor 105, pulsating pressure sensor 107, and acceleration sensor 112 to the host computer as input data for the analysis software in the host computer. The output data of the angle sensor 109 is used by the PLC 205 to determine whether the current deflection angle of the control surface 108 is consistent with the angle required by the experimental parameters. If they are inconsistent, the PLC 205 controls the servo motor 110 to drive the control surface 108 to rotate to the angle required by the experimental parameters.

[0023] Furthermore, the control unit 2 includes a housing 201, and an air switch 202, a filter 204, a switching power supply 203, a PLC 205, a servo driver 206, and an edge computing gateway 207 installed inside the housing 201. The input terminal of the air switch 202 is electrically connected to an external power supply, and the output terminal of the air switch 202 is electrically connected to the switching power supply 203. The power output terminal of the switching power supply 203 is electrically connected to the PLC 205 and the edge computing gateway 207, and the input terminal of the PLC 205 is electrically connected to the filter 204. The input terminal of the filter 204 is electrically connected to the linear displacement sensor 102, the torque strain gauge 103, the flight control sensor 105, the pulsating pressure sensor 107, and the angle sensor 109, respectively. The output terminal of the PLC 205 is electrically connected to the servo driver 206 and the edge computing gateway 207, respectively. The servo driver 206 is electrically connected to the servo motor 110 of the wing segment model 1, and the edge computing gateway 207 is electrically connected to the host computer. The enclosure 201 consists of a top cover and a bottom cover connected by hinges. The hinged connection allows for easy opening of the top cover, enabling debugging, maintenance, and upkeep of the electrical components inside the enclosure 201. The air switch 202 provides electrical protection for the electrical components within the control box, offering short-circuit and overcurrent protection. The filter 204 separates and suppresses interference in the circuit, ensuring accurate signal acquisition from the sensors. The switching power supply 203 converts AC to DC and outputs a stable voltage, providing power to the PLC 205 and the edge computing gateway 207. The PLC 205 is the core component of the control unit 2, responsible for collecting signals from various sensors and issuing control commands. The generator provides pulse signals to the servo driver 206 and signals to the edge computing gateway 207 and the host computer. The servo driver 206 is electrically connected to the servo motor 110, and controls the start and stop of the servo motor 110 to achieve the deflection of the control surface 108. The edge computing gateway 207 is used for data acquisition, download, and upload. It can temporarily store data in the event that the data cannot be transmitted to the host computer due to network abnormalities to avoid data loss. In addition, the edge computing gateway 207 can realize the rapid analysis and processing of data at the model 1 end, saving the upload of useless data to the host computer and reducing the mathematical processing pressure of the host computer.

[0024] The above are merely preferred embodiments of this utility model and are not intended to limit the scope of this utility model. Various modifications and variations can be made to this utility model by those skilled in the art. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of this utility model should be included within the protection scope of this utility model.

Claims

1. A wind tunnel flutter testing device, characterized in that, The system includes a wing segment model (1) and a control unit (2). The wing segment model (1) is fixedly connected to the wind tunnel interface by bolts. A linear displacement sensor (102) and a torque strain gauge (103) are installed on the main beam (101) of the wing segment model (1). A flight control sensor (105) is installed on the root body (104) of the wing segment model (1). A pulsating pressure sensor (107) is installed on the surface of the lower skin (106) of the wing segment model (1). The control surface (108) of the wing segment model (1) is mounted on the pivot. An angle sensor (109) is provided, and an acceleration sensor (112) is installed on the stiffener (111) of the wing segment model (1). The signal output pins of the linear displacement sensor (102), torque strain gauge (103), flight control sensor (105), pulsating pressure sensor (107) and angle sensor (109) are all electrically connected to the input pins of the control unit (2). The output pins of the control unit (2) are electrically connected to the servo motor (110) of the wing segment model (1) and the host computer, respectively.

2. The wind tunnel flutter test equipment according to claim 1, characterized in that: The control unit (2) includes a housing (201), and an air switch (202), a filter (204), a switching power supply (203), a PLC (205), a servo driver (206), and an edge computing gateway (207) installed in the housing (201). The input terminal of the air switch (202) is electrically connected to an external power supply, and the output terminal of the air switch (202) is electrically connected to the switching power supply (203). The power output terminal of the switching power supply (203) is electrically connected to the PLC (205) and the edge computing gateway (207). The input terminal of the PLC (205) is electrically connected to the filter (204); the input terminal of the filter (204) is electrically connected to the linear displacement sensor (102), the torque strain gauge (103), the flight control sensor (105), the pulsating pressure sensor (107), and the angle sensor (109), respectively; the output terminal of the PLC (205) is electrically connected to the servo driver (206) and the edge computing gateway (207), respectively. The servo driver (206) is electrically connected to the servo motor (110) of the wing segment model (1), and the edge computing gateway (207) is electrically connected to the host computer.

3. The wind tunnel flutter test equipment according to claim 1, characterized in that: The linear displacement sensor (102) is model MTI-2100.

4. The wind tunnel flutter test equipment according to claim 1, characterized in that: The number of torque strain gauges (103) is two.

5. The wind tunnel flutter test equipment according to claim 1, characterized in that: The torque strain gauge (103) is model CEA-06-250UW-350.

6. The wind tunnel flutter test equipment according to claim 1, characterized in that: The flight control sensor (105) is model ADIS16470.

7. The wind tunnel flutter test equipment according to claim 1, characterized in that: The pulsating pressure sensor (107) is model XCQ-062.

8. The wind tunnel flutter test equipment according to claim 1, characterized in that: The angle sensor (109) is model RLS RM08.

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