Aircraft aerodynamic performance test system

Abstract

The aircraft aerodynamic performance test system comprises an aircraft model, a supporting structure, an aircraft model control device and a flow field control device, the flow field control device is used for generating a flow field, the aircraft model control device is arranged outside the flow field, the aircraft model is arranged inside the flow field, and the aircraft model comprises a controller and a tilt rotor; when an aerodynamic performance test is carried out, the aircraft model is arranged on the supporting structure, and the aircraft model control device is connected with a controller of the aircraft model. A user can control the aircraft model to complete the required working condition test by controlling the aircraft model control device outside the flow field, and the efficiency of the aerodynamic performance test of the aircraft is improved.

Description

Technical Field

[0001] This application relates to the technical field of aircraft, and particularly to an aircraft aerodynamic performance test system. Background Art

[0002] In recent years, with the rapid development of the new energy field, aviation electrification has also become a trend of the times. As a new type of medium and short-distance air transportation vehicle, the electric vertical take-off and landing aircraft (abbreviated as: eVTOL) has the characteristics of electric zero-emission, high safety, high performance, etc., and has also become one of the key research directions in the new energy field.

[0003] In the research and development process of eVTOL, mastering the actual aerodynamic characteristics of the aircraft is beneficial to further optimize the aircraft configuration, and these data are usually difficult to obtain in a simulation environment.

[0004] In addition, due to the complex structure design of eVTOL, the existing aerodynamic tests of helicopters and drones, for example, cannot meet the requirements of eVTOL aerodynamic performance testing. Utility Model Content

[0005] The embodiments of this application provide an aircraft aerodynamic performance test system to solve the problems that the existing aircraft aerodynamic performance test systems of helicopters and drones are simple in control and low in efficiency and cannot meet the requirements of eVTOL wind tunnel tests.

[0006] In a first aspect, an aircraft aerodynamic performance test system is provided, including: an aircraft model, a support structure, an aircraft model control device, and a flow field control device. The flow field control device is used to control the generated flow field. The aircraft model control device is arranged outside the flow field, and the aircraft model is arranged inside the flow field. The aircraft model includes a controller and tilt rotors.

[0007] When conducting an aerodynamic performance test, the aircraft model is arranged on the support structure, and the aircraft model control device is connected to the controller of the aircraft model.

[0008] Optionally, the aircraft model further includes: a plurality of tilt servos, a plurality of control surface servos, a plurality of power motors, wings, and a tail.

[0009] The plurality of control surface servos are respectively arranged on the wings and the tail.

[0010] The plurality of tilt servos are respectively arranged in the tilt mechanisms of the tilt rotors.

[0011] The controller is a flight control computer, and the flight control computer is respectively connected to the controllers of the multiple power motors, the servo controllers of the multiple tilt rotors, and the servo controllers of the multiple control surface servos.

[0012] Optionally, each tilt rotor is used to control the tilt angle of the tilt rotor corresponding to each tilt rotor;

[0013] Each control surface servo is used to control the angle of the control surface corresponding to each control surface servo.

[0014] Optionally, the control surface servo includes at least one of the following: aileron servo, flap servo, leading edge flap servo, trailing edge flap servo, elevon servo, rudder servo, and elevator servo.

[0015] Optionally, the tilt rotor in the aircraft model and the fuselage of the aircraft model are connected in a detachable manner.

[0016] Optionally, the aircraft model is mounted upright on the support structure, or the aircraft model is mounted upside down on the support structure.

[0017] Optionally, the support structure includes a base, an arc track, and a support rod. The arc track is respectively connected to the base and the support rod; the support rod is connected to the aircraft model;

[0018] When the support rod moves on the arc track, the angle of attack of the aircraft model is changed.

[0019] Optionally, the aircraft aerodynamic performance test system further includes a balance data collector;

[0020] The balance data collector is arranged inside the aircraft model.

[0021] Optionally, the balance data collector includes a six-component full-aircraft balance and a six-component rotor balance.

[0022] Optionally, the aircraft aerodynamic performance test system further includes a parameter monitoring device;

[0023] The parameter monitoring device is connected to the flight control computer through a network cable.

[0024] This application provides an aircraft aerodynamic performance test system, including: an aircraft model, a support structure, an aircraft model control device, and a flow field control device. The flow field control device is used to generate a flow field. The aircraft model control device is arranged outside the flow field, and the aircraft model is arranged inside the flow field. The aircraft model includes a controller and a tilt-rotor. When conducting an aerodynamic performance test, the aircraft model is arranged on the support structure, and the aircraft model control device is connected to the controller of the aircraft model. Users can control the aircraft model through the aircraft model control device outside the flow field to complete the required working condition tests, improving the efficiency of the aircraft aerodynamic performance test. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] The accompanying drawings herein are incorporated into and constitute a part of this specification, showing embodiments consistent with the present application, and are used together with the specification to explain the principles of the present application.

[0026] Figure 1 Schematic diagram of the structure of the first embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0027] Figure 2 Schematic diagram of the structure of the aircraft model provided by the present application;

[0028] Figure 3 Schematic diagram of the installation method of the aircraft model provided by the present application;

[0029] Figure 4 Schematic diagram of the structure of the fourth embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0030] Figure 5 Schematic diagram of the structure of the fifth embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0031] Figure 6 Schematic diagram of the structure of the sixth embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0032] Figure 7 Schematic diagram of the structure of the seventh embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0033] Figure 8 Schematic diagram of the structure of the eighth embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0034] Figure 9 Schematic diagram of the structure of the ninth embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0035] Figure 10 Schematic diagram of the structure of the tenth embodiment of the aircraft aerodynamic performance test system provided by the present application;

[0036] Figure 11 This is a schematic structural diagram of the eleventh embodiment of the aircraft aerodynamic performance test system provided by this application.

[0037] Explanation of reference numerals: 10 - aircraft model, 20 - support structure, 30 - aircraft model control device, 40 - data acquisition device, 50 - flow field control device, 60 - parameter monitoring device, 70 - signal conditioner, 101 - controller, 1011 - flight control computer, 102 - tilt servo, 103 - flap servo, 104 - power motor, 105 - tilt rotor, 106 - wing, 107 - tail, 201 - base, 202 - arc track, 203 - strut.

[0038] Through the above-mentioned drawings, specific embodiments of this application have been shown, and there will be more detailed descriptions hereinafter. These drawings and textual descriptions are not intended to limit the scope of the concept of this application in any way, but to illustrate the concept of this application to those skilled in the art by referring to specific embodiments. Detailed implementation manners

[0039] Here, the exemplary embodiments will be described in detail, and the examples are shown in the drawings. When the following description refers to the drawings, unless otherwise indicated, the same numbers in different drawings represent the same or similar elements. The implementation manners described in the following exemplary embodiments do not represent all implementation manners consistent with this application. On the contrary, they are merely examples of devices and methods consistent with some aspects of this application as detailed in the appended claims.

[0040] First, the terms involved in this application are explained:

[0041] Electric vertical take-off and landing aircraft (abbreviation: eVTOL): Its main feature different from conventional aircraft is that it can take off and land vertically, adopts distributed propulsion, and uses an all-electric / hybrid power drive system to provide power. A tilt-rotor aircraft is a type of vertical take-off and landing aircraft that transitions between vertical take-off and landing or hovering and horizontal flight by changing the orientation of the rotors.

[0042] During the research and development of eVTOL, in order to understand the actual aerodynamic characteristics of the aircraft, wind tunnel tests are usually required to simulate the flight state of the real aircraft in the air. On the one hand, wind tunnel tests can supplement and verify the aerodynamic data obtained in the simulation environment. On the other hand, wind tunnel tests can also measure the mutual interference between the rotors and wings of the aircraft, thereby further optimizing the configuration design, manufacturing, and modification of the aircraft. These data are usually difficult to obtain in the simulation environment.

[0043] At present, some aerodynamic performance tests of aircraft for wind tunnel tests of helicopters and drones are simple and inefficient, and cannot meet the requirements of eVTOL aerodynamic performance tests.

[0044] The following uses specific embodiments to elaborate in detail on the technical solutions of the present application and how the technical solutions of the present application solve the above technical problems. These several specific embodiments below can be combined with each other, and the same or similar concepts or processes may not be repeated in some embodiments. The embodiments of the present application will be described below in conjunction with the accompanying drawings.

[0045] Embodiment 1

[0046] Figure 1 It is a schematic structural diagram of Embodiment 1 of the aircraft aerodynamic performance test system provided by the present application. As Figure 1 shown, the aircraft aerodynamic performance test system includes:

[0047] An aircraft model 10, a support structure 20, an aircraft model control device 30, and a flow field control device 50. Among them, the aircraft model 10 includes a controller 101 and tilt rotors 105. For example, the controller can be a flight control computer.

[0048] Among them, the aircraft model 10 is arranged on the support structure 20 and in the flow field generated by the flow field control device 50. The flow field can be an air flow generated by a wind tunnel. The stationary aircraft model is arranged in the wind tunnel. At this time, the flow field control device 50 is a wind tunnel control device; the flow field can also be a flow field generated by a moving device. For example, the aircraft model 10 is arranged on a moving ground belt, or on a moving vehicle / rail vehicle, and the movement drives the aircraft model to move to generate air flow. At this time, the flow field control device 50 is a device for controlling the movable device. The flow field control device 50 can adjust the flow field parameters.

[0049] Among them, the aircraft model 10 and the support structure 20 are arranged in the flow field, and the aircraft model control device 30 is outside the flow field.

[0050] When conducting an aircraft aerodynamic performance test, the aircraft model 10 is arranged on the support structure 20, and the flow field control device 50 is turned on to generate a flow field. The aircraft model control device 30 is communicatively connected to the controller 101 in the aircraft model 10, and data transmission can be carried out by means of wired cable communication, such as universal serial bus, serial port communication, CAN bus, etc.; it can also be transmitted by means of wireless communication, such as wifi local area network communication, Bluetooth communication, radio frequency communication, etc.

[0051] The aircraft model control device 30 is used to control the operation of the equipment in the aircraft model 10, such as controlling the tilt angle of the tilt rotors in the aircraft model.

[0052] Above the support structure 20, the aircraft model 10 can be fixed. When the support structure 20 changes, it can drive the angle of attack of the aircraft model 10 to change.

[0053] Optionally, the flow field control device 50 is connected to the aircraft model control device 30.

[0054] Optionally, the aircraft model control device 30 is further configured to receive the test data of the aircraft model 10.

[0055] Optionally, the aircraft aerodynamic performance test system further includes a data acquisition device 40. The data acquisition device 40 is connected to the controller 101, and the data acquisition device 40 is configured to separately receive the test data of the aircraft model 10.

[0056] Optionally, the data acquisition device 40 is integrated in the aircraft model control device 30.

[0057] The aircraft aerodynamic performance test system provided by this embodiment can, when an aerodynamic performance test is required, control the flight control computer connected thereto through the aircraft model control device outside the flow field, and then control the aircraft model to complete the required working condition test. Through this system, the user can control outside the flow field, improving the efficiency of the aerodynamic performance test of the aircraft model.

[0058] The aircraft model needs to conduct a large number of tests in the flow field. Through the above-mentioned aircraft aerodynamic performance test system, one staff member is responsible for sending instructions, and another staff member is responsible for data acquisition. A total of two personnel can achieve instruction sending and data acquisition tasks under different working conditions, reducing the number of test personnel, optimizing the test cost, and further improving the operation efficiency of the entire test process.

[0059] Embodiment 2

[0060] The configuration of the aircraft model for testing can be various. Different rotor combinations can be adopted, or different tail fin forms can be used. For example, for the rotor, a combination of tiltrotor and fixed rotor can be used, or all tiltrotors can be used; for another example, for the tail fin, a V-type tail fin can be used, or a configuration of a vertical tail fin and a horizontal tail fin can be adopted. On the basis of the above-mentioned Embodiment 1, Figure 2 This is a schematic structural diagram of the aircraft model provided by this application, as Figure 2 shown, the configuration of the aircraft model is a combination of tiltrotor and fixed rotor, and the tail fin adopts a V-type tail fin.

[0061] Among them, the controller 101 in the aircraft model is the flight control computer 1011. In addition to the flight control computer 1011, the aircraft model also includes multiple tilt servos 102, multiple control surface servos 103, power motors 104, tilt rotors 105, wings 106 and a tail wing 107;

[0062] Among them, the control surface can be the control surface of the flap on the wing, the control surface of the aileron, or the control surface of the rudder on the tail wing, the control surface of the elevator, or the control surface of the elevon. The control surface can also refer to the combined control surface of several of the above. Therefore, the control surface servo refers to the control surface servo of any one or several combinations of the above. The control surface servo is used to control the angle of the corresponding control surface.

[0063] An exemplary one is in Figure 2 where the control surfaces include the control surface of the elevon on the tail wing and the control surface of the elevator on the flap. Therefore, multiple control surface servos 103 are respectively arranged on the flap of the wing and the tail wing. The control surface servo is used to drive the corresponding control surface angle of the control surface servo. After the angle is changed, the pulling force direction of the aircraft model can be changed, and then the local attitude of the aircraft model can be adjusted.

[0064] Multiple tilt servos 102 are respectively arranged in the tilting mechanism of the tilt rotor 105. An exemplary one is in Figure 2 where the 4 rotors above the 4 tilt servos 102 are tilt rotors 105. Each tilt servo 102 is used to drive the tilting angle of the corresponding tilt rotor. After the tilting angle is changed, the pulling force direction can be changed.

[0065] Optionally, the tilt rotor 105 is connected to the fuselage of the aircraft model in a detachable manner, and the tilt rotor 105 can be disassembled or installed according to needs for corresponding aerodynamic performance tests; after the tilt rotor is installed, a powered full-aircraft aerodynamic performance test can be carried out under the action of the driving force, and a non-powered full-aircraft aerodynamic performance test can be carried out after the tilt rotor is disassembled.

[0066] The flight control computer 1011 is respectively connected to the controllers of multiple power motors 104, the servo controllers of multiple tilt servos 102 and the servo controllers of multiple control surface servos 103. The power motor 104 is used to control the rotation speed of the fixed rotor.

[0067] The aircraft aerodynamic performance test system provided in this embodiment has a flight control computer 101 inside the aircraft model connected to the controllers of multiple power motors 104, the servo controllers of multiple tilting servos 102, and the servo controllers of multiple control surface servos 103 respectively. Each servo controller is responsible for controlling the corresponding servo or motor. Through the internal connection method of the aircraft model, the flight control computer controls multiple motors to adjust the attitude of the aircraft model, and can also prevent the interference of the wire harness on the surrounding flow field of the aircraft model when connecting to the transmission system externally, thus ensuring the accuracy of the test.

[0068] Embodiment III

[0069] The configuration of the aircraft model provided in this embodiment is a configuration that entirely adopts tilt rotors and a V-shaped tail. Figure 3 As shown in the schematic diagram of the installation method of the aircraft model provided for this application, Figure 3 as shown, the support structure 20 includes a base 201, an arc track 202, and a support rod 203. The arc track 202 is connected to the base 201 and the support rod 203 respectively; the support rod 203 is connected to the aircraft model 10;

[0070] When the support rod 203 moves on the arc track 202, it can drive the aircraft model 10 to rotate, changing the angle of attack of the aircraft model.

[0071] In some embodiments, Figure 3 as shown in Figure a in Figure 3 the aircraft model is installed upright on the support rod of the support structure; as shown in Figure b in

[0072] the aircraft model is installed upside down on the support rod of the support structure. Through different installation methods, the aerodynamic performance under different working conditions can be measured.

[0073] Figure 4 As shown in the schematic diagram of the structure of Embodiment IV of the aircraft aerodynamic performance test system provided for this application, Figure 4 as shown, on the basis of Embodiment I, the aircraft model control device 30 is connected to the flight control computer 1011 through a serial port.

[0074] The cost of serial port devices and connection cables is relatively low, reducing the cost of system deployment and maintenance, especially more suitable for small and medium-sized eVTOL systems. Serial port connection can also provide high real-time performance, suitable for control and sensor data transmission that requires immediate response. Although it is not as fast as the CAN bus, for some low-rate transmission and control tasks, the serial port is sufficient.

[0075] Embodiment V

[0076] The configuration of the aircraft model provided in this embodiment is a configuration that entirely adopts tilt rotors and a V-shaped tail.Figure 5 This is a schematic structural diagram of the fourth embodiment of the aircraft aerodynamic performance test system provided by this application, as Figure 5 shown, the aircraft aerodynamic performance test system further includes a balance data collector 108;

[0077] The balance data collector 108 is arranged inside the aircraft model 10.

[0078] In some embodiments, the balance data collector 108 includes an aircraft balance and a rotor balance.

[0079] The aircraft balance can be arranged at the front, rear, middle, or side of the fuselage of the aircraft model, and the specific arrangement position is not limited.

[0080] The rotor balance can also be arranged below, above, or on the side of the rotor.

[0081] In a specific embodiment, the aircraft balance is a six-component aircraft balance 1081, and the rotor balance is a six-component rotor balance 1082;

[0082] Exemplarily, as Figure 5 shown, the six-component aircraft balance 1081 is arranged in the middle area of the fuselage of the aircraft model 10; the six-component rotor balance 1082 is arranged below the tilt-rotor. There can be multiple balance data collectors, and the number of balances can be set according to needs. One exemplary, in Figure 6 it, one six-component aircraft balance 1081 and six six-component rotor balances 1082 are arranged inside the aircraft model 10.

[0083] In some embodiments, multiple six-component rotor balances can all be arranged on one side of the wing.

[0084] In some embodiments, the balance data collector 108 is connected to the flight control computer 1011.

[0085] The six-component aircraft balance can measure the total pulling force and torque of the aircraft in various directions, so as to comprehensively evaluate its dynamic performance. By measuring the force and torque of the whole aircraft, the center of gravity position and attitude stability of the aircraft can be accurately calculated. The six-component rotor balance can accurately measure the force and torque of each rotor in various directions, help study the dynamic response and performance of the rotor in different flight stages (such as takeoff, flight, and landing), can help evaluate the force and torque distribution of the rotor in the hover state, and verify the stability and control performance during hovering. By analyzing the data collected by the rotor balance, the rotor design can be optimized, its aerodynamic characteristics can be improved, the efficiency can be increased and the vibration can be reduced, thereby improving the flight quality and passenger comfort.

[0086] Embodiment Six

[0087] Figure 6 This is a schematic structural diagram of the sixth embodiment of the aircraft aerodynamic performance test system provided by the present application, as Figure 6 shown, the aircraft aerodynamic performance test system further includes a wind tunnel data acquisition device 109;

[0088] The wind tunnel data acquisition device 109 is arranged outside the aircraft model 10, and the wind tunnel data acquisition device 109 is respectively connected to the flight control computer 1011, the balance data acquisition device 108 and the data acquisition device 40.

[0089] The wind tunnel data acquisition device 109 can summarize the data of the aircraft model in the flow field test and transmit it to the data acquisition device through a communication line.

[0090] In some embodiments, the wind tunnel data acquisition device 109 is connected to the flight control computer 1011 through a CAN bus.

[0091] The CAN bus can provide high-speed data transmission capabilities. The CAN bus supports real-time data transmission and has good synchronization performance, which is suitable for the requirements of real-time and accurate data transmission in wind tunnel tests. This is very important for the flight control computer. And using the CAN bus can reduce the complexity and cost of wiring. Compared with traditional parallel interfaces or multiple individual connections, the CAN bus only requires two wires, which can effectively reduce the number of cables and simplify the system design and installation process.

[0092] Embodiment Seven

[0093] Figure 7 This is a schematic structural diagram of the seventh embodiment of the aircraft aerodynamic performance test system provided by the present application, as Figure 7 shown, the aircraft aerodynamic performance test system further includes a parameter monitoring device 60;

[0094] The parameter monitoring device 60 is connected to the flight control computer 1011.

[0095] The parameter monitoring device 60 is arranged outside the flow field.

[0096] In some embodiments, the parameter monitoring device 60 can be integrated into the aircraft model control device 30 or the data acquisition device 40.

[0097] The parameter monitoring device 60 is mainly responsible for monitoring whether the model's own parameters meet the test requirements under different working conditions.

[0098] In some embodiments, the parameter monitoring device 60 is connected to the flight control computer 1011 through a network cable.

[0099] Ethernet connection can provide high-speed data transmission rate, which is crucial for the flight control computer, especially when dealing with a large amount of sensor data and control signals. Moreover, Ethernet connection has good stability and reliability, can work in a complex electromagnetic environment, and resist interference. Ethernet connection can support longer transmission distances.

[0100] Embodiment Eight

[0101] Figure 8 Shown in the structural schematic diagram of Embodiment Eight of the aircraft aerodynamic performance test system provided by this application, as Figure 8 shown, the aircraft aerodynamic performance test system further includes a signal conditioner 70. The signal conditioner 70 is connected to the balance data collector 108 through a power line. The signal conditioner 70 can convert a 24V voltage into a 5V voltage and can provide a 5V excitation voltage to the balance data collector 108. The signal conditioner is powered by an external 24V DC power supply.

[0102] The signal conditioner 70 can be arranged outside or inside the aircraft model.

[0103] Embodiment Nine

[0104] Figure 9 Shown in the structural schematic diagram of Embodiment Nine of the aircraft aerodynamic performance test system provided by this application, as Figure 9 shown, in the aircraft aerodynamic performance test system, a signal conditioner 70 is arranged between the wind tunnel data acquisition device 109 and the balance data collector 108.

[0105] The signal conditioner 70 is connected to the balance data collector 108 through a power line to provide a 5V excitation voltage to the balance data collector 108.

[0106] The signal conditioner 70 is connected to the balance data collector 108 through a communication line, and the signal conditioner 70 is connected to the wind tunnel data acquisition device 109 through a communication line. The signal conditioner can receive the force measurement data collected by the balance data collector through the communication line, filter and amplify it, and then transmit it to the wind tunnel data acquisition device 109 through the communication line.

[0107] Embodiment Ten

[0108] Figure 10 Shown in the structural schematic diagram of Embodiment Ten of the aircraft aerodynamic performance test system provided by this application, as Figure 10As shown in the figure, the instruction - sending PC is connected to the flight control computer through the serial port and is used to send control instructions; the parameter - monitoring PC is connected to the flight control computer through the network cable and is used to monitor the real - time parameters of the aircraft during the test (including parameters such as motor speed, tilt servo angle, elevator and rudder servo angles, wind speed, angle of attack, etc.); the wind tunnel data acquisition device is connected to the flight control computer and the six - component force sensor through CAN communication, sends the collected test data to the data acquisition PC, and the test data is exported by the staff for analysis and use.

[0109] Embodiment XI

[0110] Figure 11 As shown in the figure, it is a schematic structural diagram of Embodiment XI of the aircraft aerodynamic performance test system provided by this application. The aircraft model is an eVTOL model and is set in the wind tunnel. As Figure 11 shown, the aircraft aerodynamic performance test system mainly includes an eVTOL control system, a wind tunnel control system, a data acquisition system, and a data monitoring system.

[0111] The eVTOL control system in the control room mainly controls parameters such as the tilt angle operation amount, rudder surface angle manipulation amount, and power motor speed of the aircraft model by sending control instruction frames. The instructions are sent to the flight control computer of the aircraft model through the serial port. After receiving the instructions, the flight control controls the corresponding servos and motors to complete the instruction actions. Among them, the aircraft model is placed in the flow field environment. The instruction - sending PC sets parameters such as the angle of attack, sideslip angle, environmental wind speed, and the motor speed, tilt angle, and rudder surface angle of the aircraft according to the working conditions requirements, thereby generating a control instruction frame and sending the control instruction frame to the flight control computer inside the scaled - down model through the serial port to achieve the control of the aircraft.

[0112] The wind tunnel control system is mainly realized by the wind tunnel test center, and adjusts the angle of attack, sideslip angle, and environmental wind speed of the aircraft according to the working conditions requirements. Among them, the angle of attack is measured and calibrated manually, and the sideslip angle and environmental wind speed can be set through the control room.

[0113] The data acquisition system mainly consists of a data acquisition PC (i.e., the data acquisition device), a flight control computer, a balance data acquisition device, and a wind tunnel data acquisition device. The data acquisition PC is mainly responsible for real - time collecting wind tunnel test data through the CAN bus (including the angle of attack, sideslip angle, wind speed, motor speed, tilt angle, rudder surface angle of the aircraft from the flight control computer, and the measured torque feedback from the balance), and timely exporting the data to the staff for analysis and decision - making on the test direction.

[0114] The parameter monitoring computer is mainly responsible for monitoring whether the model's own parameters meet the test requirements under different working conditions. The monitored parameters mainly include the wind speed, angle of attack, motor speed, tilt servo angle, control surface servo angle in the environment where the aircraft is located, as well as the unlocking status information and health status of each motor, etc. Among them, the unlocking status information of the motor refers to whether the motor is locked. When a certain motor is locked, it means that the motor cannot be controlled. Users can determine whether the model can be tested normally according to the monitored unlocking status of the motor; the health status of the motor also refers to the working state and performance of the motor, which is used to describe the running health status of the motor and predict its possible failures. For example, when the temperature is too high, the health status decreases.

[0115] Finally, it should be noted that: After considering the specification and the content disclosed herein, those skilled in the art will easily think of other implementation schemes of the present utility model. The present utility model aims to cover any variations, uses or adaptive changes of the present utility model. These variations, uses or adaptive changes follow the general principles of the present utility model and include the common general knowledge or conventional technical means in the technical field not disclosed in the present utility model. It is not limited to the precise structure already described and shown in the drawings, and various modifications and changes can be made without departing from its scope. The scope of the present utility model is only limited by the appended claims.

Claims

1. An aircraft aerodynamic performance test system, characterized in that, Comprising: An aircraft model, a support structure, an aircraft model control device, and a flow field control device. The flow field control device is used to control the generated flow field. The aircraft model control device is arranged outside the flow field, and the aircraft model is arranged inside the flow field. The aircraft model includes a controller and tilt rotors. When conducting an aerodynamic performance test, the aircraft model is arranged on the support structure, and the aircraft model control device is connected to the controller of the aircraft model.

2. The aircraft aerodynamic performance test system according to claim 1, characterized in that, The aircraft model further includes: a plurality of tilt servos, a plurality of control surface servos, a plurality of power motors, a wing, and a tail. The plurality of control surface servos are respectively arranged on the wing and the tail. The plurality of tilt servos are respectively arranged in the tilt mechanisms of the tilt rotors. The controller is a flight control computer, and the flight control computer is respectively connected to the controllers of the plurality of power motors, the servo controllers of the plurality of tilt servos, and the servo controllers of the plurality of control surface servos.

3. The aircraft aerodynamic performance test system according to claim 2, wherein Each tilt servo is used to control the tilt angle of the corresponding tilt rotor of each tilt servo. Each control surface servo is used to control the angle of the corresponding control surface of each control surface servo.

4. The aircraft aerodynamic performance test system according to claim 2 or 3, characterized in that, The control surface servo includes at least one of the following: aileron servo, flap servo, leading edge flap servo, trailing edge flap servo, elevator rudder servo, rudder servo, and elevator servo.

5. The aircraft aerodynamic performance test system according to any one of claims 1 to 3, characterized in that The tilt rotors in the aircraft model and the fuselage of the aircraft model are connected in a detachable manner.

6. The aircraft aerodynamic performance test system according to any one of claims 1 to 3, characterized in that The aircraft model is installed upright on the support structure, or the aircraft model is installed upside down on the support structure.

7. The aircraft aerodynamic performance test system according to any one of claims 1 to 3, characterized in that, The support structure includes a base, an arc track, and a support rod. The arc track is respectively connected to the base and the support rod; the support rod is connected to the aircraft model. When the support rod moves on the arc track, the angle of attack of the aircraft model is changed.

8. The aircraft aerodynamic performance test system according to any one of claims 1 to 3, characterized in that, The aircraft aerodynamic performance test system further includes a balance data collector. The balance data collector is arranged inside the aircraft model.

9. The aircraft aerodynamic performance test system according to claim 8, characterized in that, The balance data collector includes a six-component full-aircraft balance and a six-component rotor balance.

10. The aircraft aerodynamic performance test system according to claim 2 or 3, characterized in that, The aircraft aerodynamic performance test system further includes a parameter monitoring device. The parameter monitoring device is connected to the flight control computer through a network cable.

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