Aircraft balance aerodynamic force testing device and vehicle-mounted testing device thereof

Abstract

The invention relates to an aircraft balance aerodynamic force testing device and a vehicle-mounted testing device thereof. The aircraft balance aerodynamic force testing device comprises a fixing part, a bearing part, a force sensor group connected between the bearing part and the fixing part, and a controller connected with the force sensor group. The fixing part and the bearing part are arranged at intervals in the vertical direction, the fixing part is used for being installed on a mobile device, and the bearing part is used for bearing a to-be-tested aircraft; the force sensor group comprises at least one force sensor sub-group, the force sensor sub-group is configured to collect at least one load electric signal of the to-be-tested aircraft acting on the bearing part in the vertical direction, the lateral direction or the course direction, and one force sensor sub-group is configured to collect the load electric signal in the same direction; the controller is configured to receive the load electric signal and solve the aerodynamic parameters of the aircraft to be tested according to the load electric signal. According to the scheme provided by the invention, the technical problem that the size of a large-scale aircraft model or a 1: 1 real aircraft is limited under a wind tunnel aerodynamic force test can be solved.

Description

Technical Field

[0001] The present application relates to the technical field of aircraft testing, and in particular to an aircraft balance aerodynamic test device and a vehicle-mounted test device thereof. Background Art

[0002] In aircraft aerodynamic testing, wind tunnel tests are often used to obtain aircraft aerodynamic parameters. However, the rental cost of wind tunnel tests is high, and the wind tunnel tests themselves are also limited by size. Therefore, how to conduct aerodynamic testing on large-scale aircraft models or 1:1 real aircraft needs to be further solved. Summary of the invention

[0003] In order to solve or partially solve the problems existing in the related art, the present application provides an aircraft balance aerodynamic test device and a vehicle-mounted test device thereof.

[0004] In a first aspect, the present application provides an aircraft balance aerodynamic force testing device, comprising a fixing portion, a bearing portion, a force sensor group connected between the bearing portion and the fixing portion, and a controller connected to the force sensor group;

[0005] The fixing portion and the carrying portion are arranged at intervals in a vertical direction, wherein the fixing portion is used to be installed on a mobile device, and the carrying portion is used to carry an aircraft to be tested;

[0006] The force sensor group includes at least one force sensor subgroup, and the force sensor subgroup is configured to collect the load electrical signal of at least one of the vertical, lateral or heading directions of the aircraft to be tested acting on the load-bearing part, wherein one of the force sensor subgroups is configured to collect the load electrical signal in the same direction;

[0007] The controller is configured to receive the load electrical signal and solve the aerodynamic parameters of the aircraft to be tested according to the load electrical signal.

[0008] The aircraft balance aerodynamic test device provided in the embodiment of the present application is applied to vehicle-mounted testing. The aircraft balance aerodynamic test device includes a fixing part, a load-bearing part, a force sensor group and a controller. The aircraft to be tested can be installed on the load-bearing part, and the fixing part can be installed on the mobile device. The force sensor group connects the load-bearing part and the fixing part and is connected to the controller. When the mobile device moves, the aircraft to be tested is driven to move together to simulate the flight of the aircraft to be tested, and during the movement, at least one force sensor subgroup in the force sensor group collects a load electrical signal of at least one of the vertical, lateral or heading directions of the aircraft to be tested acting on the load-bearing part, and one force sensor subgroup is configured to collect load electrical signals in the same direction. Finally, after receiving the load electrical signals collected by the force sensor group, the controller can solve the aerodynamic parameters of the aircraft to be tested according to the load electrical signals. The embodiment of the present application has a simple structure, can measure the aerodynamic parameters of an aircraft, and provides a vehicle-mounted testing solution that is different from a wind tunnel test, thereby solving the technical problem of size limitations of large-scale aircraft models or 1:1 real aircraft under wind tunnel aerodynamic testing, and further effectively verifying whether the aerodynamic force of the aircraft to be tested meets the requirements.

[0009] In an optional embodiment, the force sensor group includes at least one force sensor subgroup, and one force sensor subgroup includes at least one force sensor. If the number of the force sensors is more than two, all the force sensors in the force sensor group are arranged in the same plane.

[0010] In an optional embodiment, the force sensor subgroup includes a first force sensor, which is connected between the load-bearing portion and the fixing portion in a vertical direction and is used to collect a vertical load acting on the load-bearing portion by the aircraft to be tested; and / or,

[0011] The force sensor subgroup includes a second force sensor, which is connected between the load-bearing portion and the fixing portion along a first transverse direction and is used to collect a lateral load of the aircraft to be tested acting on the load-bearing portion; and / or,

[0012] The force sensor subgroup includes a third force sensor, which is connected between the load-bearing portion and the fixing portion along a second transverse direction and is used to collect a heading load of the aircraft to be tested acting on the load-bearing portion.

[0013] In an optional embodiment, four first force sensors are provided, and the four first force sensors are arranged corresponding to the four corners of the virtual rectangle where the bearing part is located, and are connected between the bearing part and the fixing part in the vertical direction at positions corresponding to the four corners; and / or,

[0014] There are two second force sensors, which are distributed in the virtual rectangle where the bearing portion is located and symmetrically distributed about the center point of the virtual rectangle, and are arranged in opposite directions along the first transverse direction; and / or,

[0015] The third force sensor is provided with one, and the third force sensor is provided along the center line of the virtual rectangle where the bearing portion is located, and the extension direction of the center line is the same as the heading of the aircraft to be tested.

[0016] In an optional embodiment, the vertical load includes lift, rolling moment and pitching moment, and the controller is configured as follows:

[0017] summing the four voltage signals collected by the first force sensor to obtain a signal output of lift; and / or,

[0018] Subtracting the sum of the two left voltage signals and the sum of the two right voltage signals collected by the first force sensor to obtain a signal output of the rolling moment; and / or,

[0019] The signal output of the pitching moment is obtained by calculating the difference between the sum of the two front-side voltage signals and the sum of the two rear-side voltage signals collected by the first force sensor.

[0020] In an optional embodiment, the lateral load includes a lateral force and a yaw moment, and the controller is configured as follows:

[0021] The difference between the front side voltage signal and the rear side voltage signal collected by the second force sensor is calculated to obtain a signal output of the lateral force; and / or,

[0022] The inverse of the front voltage signal and the inverse of the rear voltage signal collected by the second force sensor are summed to obtain a signal output of the yaw moment.

[0023] In an optional embodiment, the heading load includes resistance, and the controller is configured to output the voltage signal collected by the third force sensor as a signal of the resistance.

[0024] In an optional implementation, the force sensor is an S-type force sensor, and / or the force sensor is connected between the bearing portion and the fixing portion via a fisheye spherical bearing.

[0025] A second aspect of the present application provides a vehicle-mounted aircraft balance aerodynamic force testing device, comprising a mobile device and the aircraft balance aerodynamic force testing device as described above;

[0026] The fixed part of the aircraft balance aerodynamic test device is installed on the mobile device and is used to move with the mobile device to measure the aerodynamic parameters of the aircraft to be tested on the supporting part of the aircraft balance aerodynamic test device during the movement of the mobile device.

[0027] In an optional embodiment, it also includes:

[0028] an adjusting mechanism provided on the carrying portion, the adjusting mechanism being configured to carry the aircraft to be tested and adjust the pitch angle of the aircraft to be tested; and / or,

[0029] A bracket connected between the fixing portion and the moving device, the bracket being used to support the aircraft balance aerodynamic test device to a target height.

[0030] It should be understood that the foregoing general description and the following detailed description are exemplary and explanatory only and are not restrictive of the present application. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] The above and other objects, features and advantages of the present application will become more apparent by describing in more detail the exemplary embodiments of the present application in conjunction with the accompanying drawings, wherein the same reference numerals generally represent the same components in the exemplary embodiments of the present application.

[0032] Figure 1 It is a structural schematic diagram of an aircraft balance aerodynamic force test device shown in the present application;

[0033] Figure 2 yes Figure 1 A top view with the bearing portion removed;

[0034] Figure 3 It is a structural schematic diagram of the vehicle-mounted aircraft balance aerodynamic force test device shown in the present application;

[0035] In the figure: 1. Aircraft balance aerodynamic test device; 10. First force sensor; 11. Force sensor No. 1; 12. Force sensor No. 2; 13. Force sensor No. 3; 14. Force sensor No. 4; 20. Second force sensor; 21. Force sensor No. 5; 22. Force sensor No. 6; 30. Third force sensor; 31. Force sensor No. 7; 40. Fixing part; 50. Bearing part; 6. Adjusting mechanism; 7. Bracket; 8. Moving device. DETAILED DESCRIPTION

[0036] The embodiments of the present application will be described in more detail below with reference to the accompanying drawings. Although the embodiments of the present application are shown in the accompanying drawings, it should be understood that the present application can be implemented in various forms and should not be limited by the embodiments described herein. On the contrary, these embodiments are provided to make the present application more thorough and complete, and to fully convey the scope of the present application to those skilled in the art.

[0037] The terms used in this application are for the purpose of describing specific embodiments only and are not intended to limit this application. The singular forms of "a", "said" and "the" used in this application and the appended claims are also intended to include plural forms unless the context clearly indicates other meanings. It should also be understood that the term "and / or" used herein refers to and includes any or all possible combinations of one or more associated listed items.

[0038] It should be understood that although the terms "first", "second", "third", etc. may be used in this application to describe various information, this information should not be limited to these terms. These terms are only used to distinguish the same type of information from each other. For example, without departing from the scope of this application, the first information may also be referred to as the second information, and similarly, the second information may also be referred to as the first information. Thus, the features defined as "first" and "second" may explicitly or implicitly include one or more of the features. In the description of this application, the meaning of "multiple" is two or more, unless otherwise clearly and specifically defined.

[0039] In the related art, in aircraft aerodynamic testing, aircraft aerodynamic parameters are often obtained through wind tunnel testing. However, the rental cost of wind tunnel testing is high, and the wind tunnel testing itself is also limited by size. Therefore, how to conduct aerodynamic testing on large-scale aircraft models or 1:1 real aircraft needs to be further solved.

[0040] In response to the above problems, the present application provides an aircraft balance aerodynamic test device with a simple structure, which can measure the aerodynamic parameters of the aircraft and provide a vehicle-mounted test solution different from the wind tunnel test, thereby solving the technical problem of size limitations of large-scale aircraft models or 1:1 real aircraft under wind tunnel aerodynamic testing, and effectively verifying whether the aerodynamic force of the aircraft to be tested meets the requirements.

[0041] The technical solution of the present application is described in detail below with reference to the accompanying drawings.

[0042] Figure 1 It is a structural schematic diagram of the aircraft balance aerodynamic test device shown in the present application.

[0043] See also Figure 1The present application shows an aircraft balance aerodynamic test device, which includes a fixing portion 40, a bearing portion 50, a force sensor group connected between the bearing portion 50 and the fixing portion 40, and a controller connected to the force sensor group.

[0044] like Figure 2 As shown, the fixing portion 40 and the carrying portion 50 are arranged at intervals in the vertical direction, wherein the fixing portion 40 is used to be installed on the mobile equipment 8, and the carrying portion 50 is used to carry the aircraft to be tested.

[0045] The aircraft balance aerodynamic test device 1 is applied to vehicle-mounted testing. The fixing part 40 is installed on a mobile device 8, which can be a truck, and the carrying part 50 carries the aircraft to be tested. Figure 1 As shown, the fixing part 40 and the bearing part 50 can be two rectangular plates with similar structures, one of which is fixed on the mobile device 8 as a bottom plate, and the mobile device 8 can be a truck, and the other is regarded as a loading platform to carry the aircraft to be tested. When the truck carries the aircraft to be tested on a wide and flat road for on-board testing, the aerodynamic parameters of the aircraft to be tested are measured by the aircraft balance aerodynamic test device, thereby solving the technical problem of size limitation in wind tunnel aerodynamic testing.

[0046] The force sensor group includes at least one force sensor subgroup, which is configured to collect electrical load signals of at least one of the vertical, lateral or heading directions of the aircraft to be tested acting on the load-bearing part 50, wherein one force sensor subgroup is configured to collect electrical load signals in the same direction.

[0047] Among them, the aerodynamic parameters mainly include six components of the aircraft to be tested, namely three axial forces (drag Fx, side force Fy, lift Fz) and three moments (rolling moment Mx, pitching moment My, yaw moment Mz). Among them, lift Fz, rolling moment Mx, pitching moment My are vertical loads, side force Fy, yaw moment Mz are lateral loads, and drag Fx is heading load. The force sensor group is grouped according to the measured load direction, and the number of force sensor subgroups required and the measured load direction can be determined according to actual needs.

[0048] It should be noted that if the aerodynamic parameter to be measured is the heading load, then one force sensor can be used to achieve it, while if the aerodynamic parameter to be measured is the load in other directions, then multiple force sensors are required to achieve it together. Therefore, the force sensor group includes at least one force sensor subgroup, and a force sensor subgroup includes at least one force sensor. If the number of force sensors is more than two, all force sensors in the force sensor group are arranged in the same plane. When all force sensors are arranged in the same plane, there is no need to convert and calculate the force arms between the force sensors, thereby simplifying the measurement of the components in the embodiment of the present application.

[0049] In practice, the three axial forces and three moments of the aircraft to be tested need to be measured. Therefore, this embodiment takes the measurement of loads in three directions as an example, and the force sensor group includes a first force sensor subgroup, a second force sensor subgroup and a third force sensor subgroup. The first force sensor subgroup is used to collect the vertical load of the aircraft to be tested acting on the load-bearing part 50, the second force sensor subgroup is used to collect the lateral load of the aircraft to be tested acting on the load-bearing part 50, and the third force sensor subgroup is used to collect the heading load of the aircraft to be tested acting on the load-bearing part 50.

[0050] The controller is configured to receive the load electrical signal and solve the aerodynamic parameters of the aircraft to be tested according to the load electrical signal.

[0051] Among them, the aerodynamic parameters collected by the force sensor group exist in the form of load electrical signals, specifically voltage signals, and data processing and conversion are required. The loading platform and the additional devices above it still have a certain weight. Therefore, if the controller directly solves the aerodynamic parameters of the received load electrical signals, then the load electrical signals received here can be regarded as obtained after peeling according to the actual situation. How to peel is a delicate matter for those skilled in the art, and will not be described in detail here. After receiving the load electrical signal, the controller first determines the signal output of each component measured by the aircraft balance aerodynamic test device 1 according to the voltage signals collected by each force sensor. Thereafter, six components are solved according to the signal output of each component measured and the set correction coefficient.

[0052] The aircraft balance aerodynamic test device 1 provided in the embodiment of the present application is used in vehicle-mounted testing. When the mobile device 8 drives the aircraft to be measured to move, the force sensor group will collect one or more voltage signals corresponding to the vertical, lateral and heading directions of the aircraft to be tested acting on the supporting part 50. Calculation of one or more voltage signals in the same group can obtain signal outputs of multiple components, thereby obtaining the aerodynamic parameters of the aircraft to be tested.

[0053] The following will further describe the aircraft balance aerodynamic test device of the present application. Figure 1 and Figure 2 The present application shows an aircraft balance aerodynamic test device 1, which includes a fixed part 40, a load-bearing part 50, a force sensor group connected between the load-bearing part 50 and the fixed part 40, and a controller connected to the force sensor group, wherein the force sensor is an S-type force sensor, and the force sensor is connected between the load-bearing part 50 and the fixed part 40 through a fisheye joint bearing.

[0054] The fixing portion 40 and the carrying portion 50 are arranged at intervals in the vertical direction, wherein the fixing portion 40 is used to be installed on the mobile device 8 , and the carrying portion 50 is used to carry the aircraft to be tested.

[0055] The force sensor group includes three force sensor subgroups, which are configured to collect electrical signals of loads in the vertical, lateral and directional directions of the aircraft to be tested acting on the load-bearing part 50, and one force sensor subgroup is configured to collect electrical signals of loads in the same direction. The first force sensor subgroup collects electrical signals of loads in the vertical direction of the aircraft to be tested acting on the load-bearing part 50, and the force sensors therein include four first force sensors 10; the second force sensor subgroup collects electrical signals of loads in the lateral direction of the aircraft to be tested acting on the load-bearing part 50, and the force sensors therein include two second force sensors 20; the third force sensor subgroup collects electrical signals of loads in the directional direction of the aircraft to be tested acting on the load-bearing part 50, and the force sensors therein include one third force sensor 30. The seven force sensors in this embodiment are all S-type force sensors and are distributed at multiple points on the same plane. They have the characteristics of large load and good rigidity. All S-type force sensors are movably connected to the fixed part 40 / the load-bearing part 50 using fisheye joint bearings, so that the force sensors have the freedom to rotate around the three coordinate axes, which can minimize the interference between the various components and the accuracy can reach 0.1% FS (Full Scale) level.

[0056] In the embodiment of the present application, the vertical, lateral and heading loads are collected and acquired by the first force sensor 10, the second force sensor 20 and the third force sensor 30 respectively. For the convenience of distinction, the positive directions of the vertical, lateral and heading directions are defined to correspond to the positive directions of the Z axis, the Y axis and the X axis respectively. The X axis is defined as the fuselage direction of the aircraft to be tested. For the aircraft, the direction from the tail of the aircraft to the nose is the positive direction of the X axis; the Y axis is defined as the wing direction of the aircraft to be tested, and the direction from the left wing to the right wing is the positive direction of the Y axis; the Z axis is defined as the vertical axis of the aircraft to be tested, perpendicular to the plane determined by the X axis and the Y axis, and the positive direction of the Z axis is from the bottom of the aircraft to the top.

[0057] In some embodiments, the force sensor subgroup includes a first force sensor 10 , which is vertically connected between the load-bearing portion 50 and the fixing portion 40 , and is used to collect the vertical load of the aircraft to be tested acting on the load-bearing portion 50 .

[0058] Furthermore, four first force sensors 10 are provided, and the four first force sensors 10 are arranged corresponding to the four corners of the virtual rectangle where the carrying part 50 is located, and are connected between the carrying part 50 and the fixing part 40 along the vertical direction at the positions corresponding to the four corners.

[0059] like Figure 1 and Figure 2As shown, the first force sensor 10 has force sensor No. 11, force sensor No. 2 12, force sensor No. 3 13 and force sensor No. 4 14, force sensor No. 11 is in the left front position, force sensor No. 2 12 is in the right front position, force sensor No. 3 13 is in the left rear position, and force sensor No. 4 14 is in the right rear position, wherein the front, back, left and right positions can refer to the front, back, left and right positions in the above-mentioned XYZ axis. Force sensor No. 11, force sensor No. 2, force sensor No. 3, force sensor No. 13, and force sensor No. 4, force sensor No. 14 are arranged between the upper and lower fixed parts 40 and the bearing parts 50, and are connected to the fixed parts 40 and the bearing parts 50 to be connected through fisheye joint bearings, and force sensor No. 11, force sensor No. 2, force sensor No. 12, force sensor No. 3, force sensor No. 13, and force sensor No. 4, force sensor 14 are all distributed along the Z axis, and force sensor No. 11, force sensor No. 2, force sensor No. 12, force sensor No. 3, force sensor 13, and force sensor No. 4, force sensor 14 are on the same plane, then, according to the principle of balance force decomposition, it can be determined that force sensor No. 1 The sum of the four voltage signals collected by force sensor No. 11, force sensor No. 2 12, force sensor No. 3 13 and force sensor No. 4 14 can characterize the lift Fz without any additional calculation. The difference between the sum of the two voltage signals collected by force sensor No. 11 and force sensor No. 2 12 and the sum of the two voltage signals collected by force sensor No. 3 13 and force sensor No. 4 14 can characterize the rolling moment Mx. The difference between the sum of the two voltage signals collected by force sensor No. 1 1 and force sensor No. 3 13 and the sum of the two voltage signals collected by force sensor No. 2 12 and force sensor No. 4 14 can characterize the pitching moment My.

[0060] In some embodiments, the force sensor subgroup includes a second force sensor 20 connected between the load-bearing portion 50 and the fixing portion 40 along a first transverse direction, and used to collect the lateral load of the aircraft to be tested acting on the load-bearing portion 50 .

[0061] Further, two second force sensors 20 are provided, and the two second force sensors 20 are distributed in the virtual rectangle where the carrying portion 50 is located and are symmetrically distributed about the center point of the virtual rectangle, and are arranged in opposite directions along the first transverse direction.

[0062] like Figure 1 and Figure 2As shown, the second force sensor 20 has force sensor No. 5 21 and force sensor No. 6 22. Force sensor No. 5 21 is distributed in the front, and force sensor No. 6 22 is distributed in the back, and both force sensors are distributed along the Y-axis direction, and are distributed in the virtual rectangle where the bearing part 50 is located and are symmetrically distributed about the center point of the virtual rectangle. Force sensor No. 5 21 and sensor No. 6 are arranged between the fixed part 40 and the bearing part 50 distributed up and down, and are connected to the fixed part 40 and the bearing part 50 to be connected through a fisheye joint bearing, and the plane where force sensor No. 5 21 and force sensor No. 6 22 are located is the same as the plane where force sensor No. 1 1, force sensor No. 2 12, force sensor No. 3 13 and force sensor No. 4 14 are located. It should be explained that the coplanar arrangement of the force sensors can be understood as the coplanarity of the equivalent points of the force sensors. According to the balance force decomposition principle, it can be determined that the difference between the two voltage signals collected by force sensor No. 5 21 and force sensor No. 6 22 can represent the lateral force Fy, and the inverse of the sum of the two voltage signals collected by force sensor No. 5 21 and force sensor No. 6 22 can represent the yaw moment Mz.

[0063] In some embodiments, the force sensor subgroup includes a third force sensor 30 connected between the load-bearing portion 50 and the fixing portion 40 along the second transverse direction, and used to collect the heading load of the aircraft to be tested acting on the load-bearing portion 50 .

[0064] Furthermore, one third force sensor 30 is provided, and one third force sensor 30 is provided along the center line of the virtual rectangle where the carrying portion 50 is located, and the extension direction of the center line is the same as the heading of the aircraft to be tested.

[0065] like Figure 1 and Figure 2 As shown, the third force sensor 30 has a No. 7 force sensor 31. The No. 7 force sensor 31 is arranged along the center line of the virtual rectangle where the bearing part 50 is located, and the extension direction of the center line is the same as the Z-axis direction of the aircraft to be tested, wherein the specific position of the No. 7 force sensor 31 is not strictly required. The No. 7 sensor is arranged between the upper and lower fixed parts 40 and the bearing part 50, and is connected to the fixed part 40 and the bearing part 50 to be connected through a fisheye joint bearing, and the plane where the No. 7 force sensor 31 is located is also the same as the plane where the No. 5 force sensor 21 and the No. 6 force sensor 22 are located. According to the principle of balance force decomposition, it can be determined that a voltage signal collected by the No. 7 force sensor 31 can represent the resistance Fx.

[0066] Therefore, the six components of the aircraft to be tested, namely, the drag Fx, the side force Fy, the lift Fz, the rolling moment Mx, the pitching moment My, and the yaw moment Mz, can be obtained by three force sensor subgroups, wherein each force sensor subgroup collects the corresponding vertical, lateral, and heading load electrical signals. In the embodiment of the present application, a total of seven force sensors are arranged, wherein force sensors No. 1 to No. 4 are used to measure the vertical load (lift Fz, rolling moment Mx, pitching moment My), force sensors No. 5 to No. 6 are used to measure the lateral load (side force Fy, yaw moment Mz), and force sensor No. 7 is used to measure the heading load (drag Fx).

[0067] The controller is configured to receive the load electrical signal and solve the aerodynamic parameters of the aircraft to be tested according to the load electrical signal.

[0068] In order to facilitate distinction, the voltage signals collected by force sensors No. 1 to No. 7 are recorded as U1~U7, and the signal outputs of the six components of drag Fx, side force Fy, lift Fz, rolling moment Mx, pitching moment My, and yaw moment Mz are recorded as U Fx , U Fy , U Fz , U Mx , U My and U Mz .

[0069] In some embodiments, the vertical load includes lift Fz, rolling moment Mx, and pitching moment My. The controller is configured to: sum the four voltage signals collected by the first force sensor 10 to obtain a signal output of lift; and / or, subtract the sum of the two left voltage signals collected by the first force sensor 10 from the sum of the two right voltage signals to obtain a signal output of rolling moment; and / or, subtract the sum of the two front voltage signals collected by the first force sensor 10 from the sum of the two rear voltage signals to obtain a signal output of pitching moment.

[0070] In this embodiment, the first force sensor 10 is the force sensor No. 1 11, the force sensor No. 2 12, the force sensor No. 3 13 and the force sensor No. 4 14 mentioned above, then, the signal output U of the lift Fz Fz =U1+U2+U3+U4, the signal output U of the rolling moment Mx Mx =U1-U2+U3-U4, the signal output U of the pitch moment My My =U1+U2-U3-U4.

[0071] In some embodiments, the lateral load includes a lateral force and a yaw moment, and the controller is configured to: calculate the difference between the front voltage signal and the rear voltage signal collected by the second force sensor 20 to obtain a signal output of the lateral force; and / or calculate the sum of the inverse of the front voltage signal and the inverse of the rear voltage signal collected by the second force sensor 20 to obtain a signal output of the yaw moment.

[0072] In this embodiment, the second force sensor 20 is the force sensor No. 5 21 and the force sensor No. 6 22 mentioned above. Then, the signal output U of the lateral force Fy is Fy =U5-U6, signal output U of yaw moment Mz Mz =-U5-U6.

[0073] In some embodiments, the heading load includes resistance, and the controller is configured to output the voltage signal collected by the third force sensor 30 as a signal of the resistance.

[0074] In this embodiment, the third force sensor 30 is the No. 7 force sensor 31 mentioned above, then the signal output U of the resistance Fx is Fx =U7.

[0075] In the controller, after the signal output of the six components is determined, the signal output of the six components can be processed by the set correction coefficient to obtain the six components, wherein the correction coefficient can be determined based on experience, and the correction coefficient includes the main term coefficient, the first-order interference correction coefficient and the correction coefficient. The mathematical formula for solving the six components is:

[0076]

[0077] Where, Fx, Fy…Mz are the six components (unit: N or N·m) obtained by the aircraft balance aerodynamic test device 1, which represent the drag, side force, lift, rolling moment, pitching moment, and yaw moment respectively; U Fx , U Fy …U Mz The signal outputs corresponding to the six components of the aircraft balance aerodynamic test device 1; is the principal coefficient of the aircraft balance aerodynamic test device 1; is the first-order interference correction coefficient of the aircraft balance aerodynamic test device 1; c Fx 、c Fy …c Mz is the correction factor of the aircraft balance aerodynamic test device 1.

[0078] Corresponding to the above-mentioned embodiments of the aircraft balance aerodynamic test device, the present application also provides a vehicle-mounted aircraft balance aerodynamic test device and corresponding embodiments.

[0079] like Figure 3 As shown, an embodiment of the present application provides a vehicle-mounted aircraft balance aerodynamic test device, comprising a mobile device 8 and the aircraft balance aerodynamic test device 1 as described above; a fixed portion 40 of the aircraft balance aerodynamic test device 1 is installed on the mobile device 8, and is used to move with the mobile device 8, so as to measure the aerodynamic parameters of the aircraft to be tested on the bearing portion 50 of the aircraft balance aerodynamic test device 1 during the movement of the mobile device 8.

[0080] In the vehicle-mounted aircraft balance aerodynamic test device, the aircraft balance aerodynamic test device 1 can be used for vehicle-mounted testing. The fixing part 40 is installed on the mobile device 8, which can be a truck, and the carrying part 50 carries the aircraft to be tested. Figure 1 As shown, the fixing part 40 and the bearing part 50 can be two rectangular plates with similar structures, one of which is fixed on the mobile device 8 as a bottom plate, and the mobile device 8 can be a truck, and the other is regarded as a loading platform to carry the aircraft to be tested. When the truck carries the aircraft to be tested on a wide and flat road for on-board testing, the aerodynamic parameters of the aircraft to be tested are measured by the aircraft balance aerodynamic test device, thereby solving the technical problem of size limitation in wind tunnel aerodynamic testing.

[0081] In some embodiments, it also includes:

[0082] The adjusting mechanism 6 is disposed on the carrying portion 50 , and is configured to carry the aircraft to be tested and adjust the pitch angle of the aircraft to be tested.

[0083] It is understandable that the aircraft to be tested also has a certain pitch angle when it moves with the mobile device 8, which can be achieved by setting an adjustment mechanism 6 on the bearing part 50, wherein the main function of the adjustment mechanism 6 is to adjust the pitch angle of the aircraft to be tested, and one end of the adjustment mechanism 6 is connected to the aircraft to be tested, and the other end is fixed to the bearing part 50 of the aircraft balance aerodynamic test device 1. The specific structure of the adjustment mechanism 6 can be any one of the prior art, and will not be described in detail here.

[0084] In some embodiments, a bracket 7 is connected between the fixing portion 40 and the mobile device 8, and the bracket 7 is used to support the aircraft balance aerodynamic test device 1 to the target height. Generally speaking, the aircraft to be tested is relatively large in size and may not be directly matched with the mobile device 8. Therefore, a rigid bracket 7 can be used as a component to support the aircraft balance aerodynamic test device 1, the adjustment mechanism 6 and the aircraft to be tested, so that the aircraft to be tested is higher than the mobile device 8, thereby simulating the flight environment of the aircraft to be tested. One end of the bracket 7 is connected to the aircraft balance aerodynamic test device 1, and the other end is connected to the mobile device 8.

[0085] Regarding the vehicle-mounted aircraft balance aerodynamic test device in the above embodiment, the specific manner in which each module performs operations has been described in detail in the embodiment of the aircraft balance aerodynamic test device, and will not be elaborated again here.

[0086] The embodiments of the present application have been described above, and the above description is exemplary, not exhaustive, and is not limited to the disclosed embodiments. Many modifications and changes will be apparent to those of ordinary skill in the art without departing from the scope and spirit of the described embodiments. The selection of terms used herein is intended to best explain the principles of the embodiments, practical applications, or improvements to the technology in the market, or to enable other persons of ordinary skill in the art to understand the embodiments disclosed herein.

Claims

1. An aircraft balance aerodynamic test device, characterized in that: It comprises a fixing part, a bearing part, a force sensor group connected between the bearing part and the fixing part, and a controller connected to the force sensor group; The fixing portion and the carrying portion are arranged at intervals in a vertical direction, wherein the fixing portion is used to be installed on a mobile device, and the carrying portion is used to carry an aircraft to be tested; The force sensor group includes at least one force sensor subgroup, and the force sensor subgroup is configured to collect the load electrical signal of at least one of the vertical, lateral or heading directions of the aircraft to be tested acting on the load-bearing part, wherein one of the force sensor subgroups is configured to collect the load electrical signal in the same direction; The controller is configured to receive the load electrical signal and solve the aerodynamic parameters of the aircraft to be tested according to the load electrical signal.

2. The device according to claim 1, characterized in that The force sensor group includes at least one force sensor subgroup, and one of the force sensor subgroups includes at least one force sensor. If the number of the force sensors is more than two, all the force sensors in the force sensor group are arranged in the same plane.

3. The device according to claim 1 or 2, characterized in that: The force sensor subgroup includes a first force sensor, which is connected between the load-bearing portion and the fixing portion in a vertical direction and is used to collect a vertical load acting on the load-bearing portion by the aircraft to be tested; and / or, The force sensor subgroup includes a second force sensor, which is connected between the load-bearing portion and the fixing portion along a first transverse direction and is used to collect a lateral load of the aircraft to be tested acting on the load-bearing portion; and / or, The force sensor subgroup includes a third force sensor, which is connected between the load-bearing portion and the fixing portion along a second transverse direction and is used to collect a heading load of the aircraft to be tested acting on the load-bearing portion.

4. The device according to claim 3, characterized in that: There are four first force sensors, which are arranged corresponding to the four corners of the virtual rectangle where the bearing part is located, and are connected between the bearing part and the fixing part in the vertical direction at positions corresponding to the four corners; and / or, There are two second force sensors, which are distributed in the virtual rectangle where the bearing portion is located and symmetrically distributed about the center point of the virtual rectangle, and are arranged in opposite directions along the first transverse direction; and / or, The third force sensor is provided with one, and the third force sensor is provided along the center line of the virtual rectangle where the bearing portion is located, and the extension direction of the center line is the same as the heading of the aircraft to be tested.

5. The device according to claim 4, characterized in that The vertical load includes lift, rolling moment and pitching moment, and the controller is configured as follows: summing the four voltage signals collected by the first force sensor to obtain a signal output of lift; and / or, Subtracting the sum of the two left voltage signals and the sum of the two right voltage signals collected by the first force sensor to obtain a signal output of a rolling moment; and / or, The signal output of the pitching moment is obtained by subtracting the sum of the two front-side voltage signals and the sum of the two rear-side voltage signals collected by the first force sensor.

6. The device according to claim 4, characterized in that The lateral load includes lateral force and yaw moment, and the controller is configured to: The difference between the front side voltage signal and the rear side voltage signal collected by the second force sensor is calculated to obtain a signal output of the lateral force; and / or, The inverse of the front voltage signal and the inverse of the rear voltage signal collected by the second force sensor are summed to obtain a signal output of the yaw moment.

7. The device according to claim 4, characterized in that The heading load includes resistance, and the controller is configured to output the voltage signal collected by the third force sensor as a signal of the resistance.

8. The device according to claim 2, characterized in that The force sensor is an S-type force sensor, and / or the force sensor is connected between the bearing part and the fixing part through a fisheye spherical bearing.

9. A vehicle-mounted aircraft balance aerodynamic test device, characterized in that: It comprises a mobile device and an aircraft balance aerodynamic test device as claimed in any one of claims 1 to 8; The fixed part of the aircraft balance aerodynamic test device is installed on the mobile device and is used to move with the mobile device to measure the aerodynamic parameters of the aircraft to be tested on the supporting part of the aircraft balance aerodynamic test device during the movement of the mobile device.

10. The device according to claim 9, characterized in that Also includes: an adjusting mechanism disposed on the carrying portion, the adjusting mechanism being configured to carry the aircraft to be tested and adjust the pitch angle of the aircraft to be tested; and / or, A bracket connected between the fixing portion and the moving device, the bracket being used to support the aircraft balance aerodynamic test device to a target height.

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