Electric rotor configuration characteristic multifunctional test bench and test system thereof

Abstract

According to the technical scheme, the electric rotor configuration characteristic multifunctional test bench is characterized in that the electric rotor configuration characteristic multifunctional test bench comprises a rack base, a rack platform and a tilting mechanism, the rack platform is rotationally connected with the rack base, the tilting mechanism is arranged on the rack platform, a tilting rotor is further connected to the tilting mechanism, and the tilting rotor is arranged on the rack base. The rack base is used for controlling rotation and pitching rotation adjustment of the rack platform; and the tilting mechanism is used for controlling the tilting rotor to simulate tilting so as to adjust the angle of the tilting rotor, simulate the aerodynamic characteristics of the aircraft in a pitching state, static or dynamic flight and simulate the attitude of the aircraft in the flight process. And the real stress condition and other physical properties of the rotor wing under different flight attitudes can be fully simulated.

Description

Technical Field

[0001] The invention relates to the technical field of rotor testing, and more particularly to a multifunctional test bench for electric rotor configuration characteristics. Background Art

[0002] At present, with the development of aircraft, the styles of aircraft have also undergone many changes. Among them, tilt-rotors have become more efficient than fixed-wing aircraft during takeoff and landing. However, since the tilt-rotors need to control the rotors to adjust the tilt angle, the performance of the tilt-rotors needs to undergo rigorous production testing to maintain the stability of the overall structure during pitch or yaw control, and to avoid problems such as uneven force and body shaking.

[0003] The invention patent with the existing announcement number CN111516903B is a tilt-rotor UAV test bench, in which the bottom profile is a cross-shaped structure, the lower end of the column is detachably arranged in the middle of the bottom profile, the upper end of the column is detachably connected to the transversely arranged top profile, one end of the tilt axis is fixed on the top of the top profile, the other end of the tilt axis is rotatably connected to two machine base fixing plates, the two machine base fixing plates are connected through a motor base, a drive motor is fixed on the surface of the motor base away from the tilt axis, a propeller is fixed on the output shaft of the drive motor, a servo is fixed on one of the machine base fixing plates, the output shaft of the servo is connected to the tilt axis through a swing assembly, and the output shaft of the servo drives the machine base fixing plate to rotate around the tilt axis.

[0004] Although the prior art discloses a tilt-rotor UAV test bench, which performs flight simulation tests on the tilt-rotor to obtain data during tilt control, the test bench structure in the prior art is relatively simple and cannot fully simulate the actual force conditions or other physical properties of the rotor when the rotor is tilted in pitch or yaw. In view of this, the present invention provides a multifunctional test bench for rotor configuration characteristics. Summary of the invention

[0005] In view of the deficiencies in the prior art, an object of the present invention is to provide a multifunctional test bench for electric rotor configuration characteristics, which can fully simulate the actual force conditions and other physical properties of the rotor in different flight attitudes.

[0006] To achieve the above object, the present invention provides the following technical solutions:

[0007] A multifunctional test bench for electric rotor configuration characteristics, comprising a bench base, a bench platform and a tilting mechanism.

[0008] The gantry platform is rotatably connected to the gantry base, the tilt mechanism is arranged on the gantry platform, a tilt rotor is further connected to the tilt mechanism, and the gantry base is used to control the rotation and pitch rotation adjustment of the gantry platform;

[0009] The tilt mechanism is used to control the simulated tilt of the tilt rotor to adjust the angle of the tilt rotor, simulate the aerodynamic characteristics of the aircraft in a pitch state, static or dynamic flight, and simulate the attitude of the aircraft during flight.

[0010] As a further improvement of the present invention, the gantry base includes a base, a pillar and an electric push rod, a rotating mechanism is connected between the base and the pillar, the rotating mechanism includes a servo motor, a planetary reducer and a worm gear mechanism, a seat body is provided on the base, a worm gear is provided in the seat body, the worm gear is connected to the pillar, the planetary reducer is provided on the seat body, the input shaft of the planetary reducer is coaxially fixedly connected to the servo motor, the output shaft of the planetary reducer is coaxially fixedly connected to the worm, and the worm is meshed with the worm gear.

[0011] As a further improvement of the present invention, the tilt mechanism includes a fixed beam, a reinforcement tube, a C-shaped beam, an electric adjustment mounting seat, a tilt servo and a servo rocker arm, the fixed beam is connected to the rotor fixed tube, the reinforcement tube is arranged on the fixed beam, and a reinforcing rib is connected between the reinforcement tube and the fixed beam, the end of the reinforcement tube away from the reinforcement tube is connected to the electric adjustment mounting seat, the electric adjustment mounting seat includes an electric adjustment mounting seat 1 and an electric adjustment mounting seat 2, the electric adjustment mounting seat 1 and the electric adjustment mounting seat 2 are used to install the electric adjustment, and the C-shaped beam is located on the side of the reinforcement tube away from the reinforcing rib, and is used to connect the reinforcement tube and the fixed beam.

[0012] As a further improvement of the present invention, a tilt servo mounting seat is also provided on the C-beam, the tilt servo is installed on the tilt servo mounting seat, the tilt servo fastening seat is used to be connected to the tilt motor mounting seat and to fix the tilt servo to the tilt servo mounting seat, and the tilt servo is also provided with a bearing seat, and the bearing seat is provided with a bearing and a clamp for fixing the bearing.

[0013] As a further improvement of the present invention, a tilt axis is arranged at the axial position of the bearing, the tilt servo is used to drive the tilt axis to rotate, the axial end of the tilt axis is connected to a servo rocker arm, the servo rocker arm comprises a servo rocker arm 1 and a servo rocker arm 2, the servo rocker arm 1 is connected to the tilt axis, the servo rocker arm 1 and the servo rocker arm 2 are respectively connected to ball head tie rods, and the ball head tie rod between the servo rocker arm 1 and the servo rocker arm 2 is respectively connected to a servo tie rod.

[0014] As a further improvement of the present invention, the tilt rotor includes a tilt connecting plate, a motor connecting piece, a motor, a hub and a rotor, the tilt connecting plate is connected to the tilt axis, a steel shaft for connecting the motor is also provided in the tilt connecting plate, a six-component balance is also connected to the motor, the steel shaft is connected to the motor connecting piece, the motor is reinforced with the connection with the tilt connecting plate through the motor connecting piece, the output end of the motor is connected to the hub, a hub bearing is provided between the hub and the rotor, the hub bearing is connected to the hub, and an adjustment ring is also provided between the hub bearing and the hub.

[0015] As a further improvement of the present invention, a propeller clamp is connected to the side of the propeller hub bearing facing away from the propeller hub, a thrust bearing is also arranged between the propeller clamp and the rotor, a gasket is arranged between the thrust bearing and the rotor, an adjustment plate is also arranged between the thrust bearing and the propeller clamp, a tie rod ball buckle is also arranged between the propeller hub and the propeller clamp, the tie rod ball buckle is connected to the propeller hub and the propeller clamp, a pitch rod is connected to the tie rod ball buckle, an end of the pitch rod away from the tie rod ball buckle is connected to a pitch rod hinge, the side of the pitch rod hinge facing away from the pitch rod is connected to the bottom of the propeller hub, a pitch rod hinge bearing is arranged at the connection between the pitch rod hinge and the propeller hub, a retaining ring is arranged between the pitch rod articulated bearing and the propeller hub, a shaft sleeve is also arranged between the pitch rod articulated bearing and the retaining ring, and the propeller hub adjusts the tilt angle of the rotor under the action of the pitch rod.

[0016] As a further improvement of the present invention, a servo for driving a pitch rod is provided on the side of the hub facing away from the motor, a fixed ring base is also provided on the motor, an activity space is formed between the fixed ring base and the motor, the hub and the propeller clamp are arranged in the activity space, a hub bearing is also provided between the hub and the fixed ring base, the hub bearing is used to reduce the friction force when the hub rotates, a hub bearing cover is provided on the side of the hub bearing facing away from the hub, a rotating ring cover is provided on the side of the hub bearing cover facing away from the hub bearing, a bearing is also provided between the rotating ring cover and the fixed ring base, and a fixed ring bearing seat for fixing the bearing.

[0017] As a further improvement of the present invention, a bearing is further arranged between the fixed ring bearing seat and the rotating ring cover plate, a rotating ring is arranged between the bearing and the fixed ring base, and a ball head for reducing the rotational friction of the rotating ring is arranged between the rotating ring and the fixed ring base.

[0018] As a further improvement of the present invention, the test system of the electric rotor configuration characteristic multifunctional test bench is applied to the electric rotor configuration characteristic multifunctional test bench:

[0019] It includes a flight dynamics module, a sensor module, a simulation model integration module, a test instance module and a data analysis module, wherein the five modules constitute a closed-loop data flow, and the flight dynamics module, the sensor module, the simulation model integration module and the test instance module all run in a real-time embedded single-chip microcomputer;

[0020] The test instance module generates an excitation signal and sends it to the flight controller to be tested;

[0021] The simulation model integration module sends the control signal of the flight controller to be tested to the flight dynamics module;

[0022] The flight dynamics module generates angular motion and linear motion information of the multi-rotor aircraft and sends it to the sensor module;

[0023] The sensor module sends the aircraft angular motion and linear motion information with noise to the flight controller;

[0024] The data analysis module receives the expected signal and the feedback signal fed back by the flight controller to be tested.

[0025] Beneficial effects of the present invention:

[0026] The multifunctional test bench for electric rotor configuration characteristics is mainly composed of columns, test benches, rotor systems, yaw control mechanisms and pitch control mechanisms. The multifunctional test bench can measure the aerodynamic characteristics of a single rotor and multiple rotors under different numbers of rotors, different front-to-back spacing, different left-to-right spacing, different height spacing, and different flight attitudes. The rotor system can realize the adjustment of cyclic pitch and total pitch, and the tilt mechanism can realize the horizontal and vertical state switching of the rotor system. The entire tilt-rotor system can realize the spatial position transformation along the crossbeam, longitudinal beam and vertical beam. The yaw angle adjustment is realized by the worm gear mechanism, and the pitch direction angle adjustment of the test bench is realized by the electric push rod.

[0027] The rotor system is mainly composed of a pitch mechanism and a tilt mechanism, which is installed on the outside and connected to the rotor installed on the top through a series of push-pull rods. There is a fixed length pull rod between the pitch rocker arm of the blade. Therefore, the blade cyclic pitch is achieved by changing the controlled tilt, and the total pitch is adjusted by moving up or down. The tilt of the rotor system is mainly achieved by the rotation of the tilt servo to change the servo rocker arm 2, and the servo rocker arm is rotated accordingly through the servo rocker arm 1 and the servo rocker arm 2. The tilt shaft rotates with the rotation of the servo rocker arm, thereby driving the tilt of the rotor system.

[0028] The multifunctional test bench adopts a modular design, which is easy to disassemble, assemble and debug. It can measure the aerodynamic characteristics of different numbers and different spatial positions of rotors according to test requirements. Each rotor system is connected to the test bench by screws, which can be easily disassembled and installed. At the same time, in addition to the pitch and tilt functions, the lateral spacing, longitudinal spacing and height position between each rotor system can be adjusted according to test requirements. BRIEF DESCRIPTION OF THE DRAWINGS

[0029] Figure 1 It is a schematic diagram of the overall three-dimensional structure of the test bench;

[0030] Figure 2 This is the exploded view of the stand base;

[0031] Figure 3 This is the exploded view of the test bench platform;

[0032] Figure 4 This is the exploded diagram of the tilting mechanism;

[0033] Figure 5 This is the exploded diagram of the rotor mechanism;

[0034] Figure 6 This is the rotor system structure diagram;

[0035] Figure 7 This is the structural diagram of the tilting mechanism;

[0036] Figure 8 It is the structure diagram of the pitch-changing mechanism;

[0037] Fig. 9 It is a flowchart of a test method for a multi-rotor aircraft control system;

[0038] Fig.10 A schematic diagram of the test results of the horizontal longitudinal channel in the position control mode provided by an embodiment.

[0039] 1. Steel shaft; 2. Tilt connecting plate; 3. Six-component balance; 4. Motor connecting piece; 5. Motor; 6. Hub; 7. Adjusting ring; 8. Hub bearing; 9. Propeller clamp; 10. Adjusting plate; 11. Thrust bearing; 12. Gasket; 13. Rotor; 14. Retaining ring; 15. Pitch rod hinge bearing; 16. Bushing; 17. Pitch rod hinge; 18. Pitch rod; 19. Rod ball buckle; 20. Hub bearing; 21. Hub bearing cover; 22. Rotating ring cover; 23. Bearing; 24. Fixed ring bearing seat; 25. Bearing; 26. Rotating ring; 27. Ball head; 28. Fixed ring base; 29. ​​Servo lower mounting seat; 30. Servo mounting seat clamp; 31. Servo rocker arm; 32. Servo; 33. Guide limit plate; 34. Servo upper mounting seat; 35. Rudder Clamp of the machine mounting seat; 36, fixed beam; 37, reinforcing rib; 38, reinforcement pipe; 39, C-beam; 40, ESC mounting seat 1; 41, ESC mounting seat 2; 42, ESC; 43, tilt servo mounting seat; 44, tilt servo; 45, tilt servo fastening seat; 46, bearing seat; 47, clamp; 48, bearing; 49, tilt shaft; 50, servo rocker arm 1; 51, servo rocker arm 2. 51; 52. Ball head pull rod; 53. Servo pull rod; 54. T-beam; 55. Support beam 1; 56. Support beam 2; 57. Rotor fixing tube; 58. Bottom connecting plate; 59. Support; 60. Bearing; 61. Bearing seat; 62. Pillar; 63. Bushing; 64. Electric push rod; 65. Worm gear mechanism; 66. Planetary reducer; 67. Servo motor; 68. Base. DETAILED DESCRIPTION

[0040] The present invention is further described in detail below in conjunction with the accompanying drawings and embodiments. The same parts are represented by the same reference numerals. It should be noted that the words "front", "rear", "left", "right", "upper" and "lower" used in the following description refer to directions in the accompanying drawings, and the words "bottom surface" and "top surface", "inner" and "outer" refer to directions toward or away from the geometric center of a specific component, respectively.

[0041] Embodiment 1:

[0042] refer to Figures 1 to 8 As shown, a specific implementation of a multifunctional test bench for electric rotor configuration characteristics of the present invention includes a test bench base, a test bench platform and a tilt mechanism. The test bench platform is rotatably connected to the test bench base, and the tilt mechanism is arranged on the test bench platform. A tilt-rotor 13 is also connected to the tilt mechanism. Under the action of the test bench base, the tilt platform is used to control the rotation and pitch rotation adjustment of the test bench platform. Under the action of the tilt mechanism, the tilt-rotor 13 is controlled to simulate the tilt, thereby adjusting the angle of the tilt-rotor 13 to simulate the aerodynamic characteristics of the aircraft in the pitch state, static or dynamic flight, and simulate the attitude of the aircraft during the flight process.

[0043] The gantry base includes a base 68, a pillar 62 and an electric push rod 64. A rotating mechanism is connected between the base 68 and the pillar 62. The rotating mechanism includes a servo motor 67, a planetary reducer 66 and a worm gear mechanism 65. A seat body is provided on the base 68, a worm gear is provided in the seat body, the worm gear is connected to the pillar 62, and the planetary reducer 66 is provided on the seat body. The input shaft of the planetary reducer 66 is coaxially fixedly connected to the servo motor 67, and the output shaft of the planetary reducer 66 is coaxially fixedly connected to the worm. The worm is meshed with the worm gear so that the output speed is reduced through the planetary reducer 66 under the drive of the servo motor 67, and the worm is driven to rotate through the output shaft of the planetary reducer 66 to control the rotation of the worm gear, thereby driving the pillar 62 to rotate and adjust, so as to adjust the rotation of the gantry platform and simulate the rotation adjustment of the aircraft.

[0044] An ear seat is extended from the outer wall of the support column 62, one end of the electric push rod 64 is rotatably connected to the ear seat, and the other end is rotatably connected to the gantry platform. A shaft sleeve 16 is provided at the connection between one end of the electric push rod 64 connected to the gantry platform, and the gantry platform is rotated and adjusted along the shaft sleeve 16 under the pushing action of the electric push rod 64, thereby realizing simulated adjustment of the pitch angle of the gantry platform.

[0045] The gantry platform is connected to the support 62, a bottom connecting plate 58 is provided at the bottom of the gantry platform, the bottom connecting plate 58 is connected to the support 59, the end of the support 62 away from the base 68 is connected to the bearing seat 46, the bearing seat 46 is provided with a bearing 23, so as to realize the rotation connection between the support 62 and the gantry platform, the gantry platform includes a T-beam 54, a supporting crossbeam 1, a supporting crossbeam 2 and a fixing tube of the rotor 13, the supporting crossbeam 1 is provided at both ends of the T-beam 54, and the supporting crossbeam 2 is arranged in sequence along the length direction of the T-beam 54. There are multiple T-beams 54, which are respectively located between the supporting beams 1 at both ends of the T-beam 54. The T-beams 54 are distributed along both sides of the supporting beams 1 and 2. The T-beams 54 on both sides are connected and fixed by the supporting beams 1 and 2. The rotor 13 fixing tube is set between the supporting beams 1 and 2, and is connected to the T-beams 54 on both sides to form an integral gantry platform. The tilt mechanism is set on the rotor 13 fixing tube, and the rotor 13 fixing tube is connected with the tilt mechanism along both sides of the T-beam 54.

[0046] The tilt mechanism includes a fixed beam 36, a reinforcement pipe 38, a C-beam 39, an electric adjustment mounting seat, a tilt servo 44 and a servo rocker arm 31. The fixed beam 36 is connected to the fixed pipe of the rotor 13. The reinforcement pipe 38 is arranged on the fixed beam 36. A reinforcing rib 37 is further connected between the reinforcement pipe 38 and the fixed beam 36. One end of the reinforcement pipe 38 away from the reinforcement pipe 38 is connected to the electric adjustment mounting seat. The electric adjustment mounting seat includes an electric adjustment mounting seat 1 40 and an electric adjustment mounting seat 2 41. The first electric adjustment mounting seat 40 and the second electric adjustment mounting seat 41 are used to install the electric adjustment 42. The C-shaped beam 39 is located on the side of the reinforcement pipe 38 away from the reinforcement rib 37 and is used to connect the reinforcement pipe 38 and the fixed beam 36. The C-shaped beam 39 is also provided with a tilt steering gear 44 mounting seat. The tilt steering gear 44 is installed on the tilt steering gear 44 mounting seat. The tilt steering gear 44 fastening seat is used to connect with the tilt motor 5 mounting seat and to fix the tilt steering gear 44 to the tilt steering gear 44 mounting seat. The tilting servo 44 is also provided with a bearing seat 46, on which a bearing 23 and a clamp 47 for fixing the bearing 23 are provided, and a tilting shaft 49 is provided at the axial position of the bearing 23. The tilting servo 44 is used to drive the tilting shaft 49 to rotate, and the shaft end of the tilting shaft 49 is connected to the servo rocker arm 31, which includes a servo rocker arm 1 50 and a servo rocker arm 2 51. The servo rocker arm 1 50 is connected to the tilting shaft 49, and the servo rocker arm A ball head pull rod 52 is connected to the servo rocker arm 1 50 and the servo rocker arm 2 51 respectively, and a servo pull rod 53 is connected to the ball head pull rod 52 between the servo rocker arm 1 50 and the servo rocker arm 2 51 respectively, so that the tilt servo 44 is controlled to rotate under the action of the electric adjustment 42, and the tilt servo 44 drives the tilt shaft 49 to rotate, so that the servo rocker arm 1 50 rotates and drives the servo rocker arm 2 51 to tilt through the servo pull rod 53, thereby realizing the control of the tilt rotor 13.

[0047] The tilt rotor 13 includes a tilt connection plate 2, a motor connection member 4, a motor 5, a hub 6 and a rotor 13, the tilt connection plate 2 is connected to the tilt shaft 49, a steel shaft 1 for connecting the motor 5 is also provided in the tilt connection plate 2, a six-component balance 3 is also connected to the motor 5, and the deflection angle of the rotor 13 can be judged by the six-component balance, wherein the steel shaft 1 is connected to the motor connection member 4, the motor 5 is reinforced with the connection with the tilt connection plate 2 through the motor connection member 4, the output end of the motor 5 is connected to the hub 6, a hub bearing 21 is provided between the hub 6 and the rotor 13, the hub bearing 21 is connected to the hub 6, an adjustment ring 7 is also provided between the hub bearing 21 and the hub 6, a propeller clamp 9 is connected to the side of the hub bearing 21 away from the hub 6, a thrust bearing 11 is also provided between the propeller clamp 9 and the rotor 13, and the thrust bearing 11 is connected to the rotor 1 3, an adjusting sheet 10 is also provided between the thrust bearing 11 and the blade clamp 9, a tie rod ball buckle 19 is also provided between the blade hub 6 and the blade clamp 9, the tie rod ball buckle 19 is connected to the blade hub 6 and the blade clamp 9, a pitch rod 18 is connected to the tie rod ball buckle 19, one end of the pitch rod 18 away from the tie rod ball buckle 19 is connected to a pitch rod hinge 17, the side of the pitch rod hinge 17 away from the pitch rod 18 is connected to the bottom of the blade hub 6, a pitch rod hinge bearing 15 is provided at the connection between the pitch rod hinge 17 and the blade hub 6, a retaining ring 14 is provided between the pitch rod hinge bearing 15 and the blade hub 6, and a shaft sleeve 16 is also provided between the pitch rod hinge bearing 15 and the retaining ring 14, so that the tilt angle of the rotor 13 can be controlled under the action of the blade hub 6 and the pitch rod 18, thereby achieving the angle adjustment of the rotor 13.

[0048] A steering gear 32 for driving the pitch-changing pull rod 18 is arranged on the side of the propeller hub 6 away from the motor 5, a fixed ring base 28 is also arranged on the motor 5, an activity space is formed between the fixed ring base 28 and the motor 5, the propeller hub 6 and the propeller clamp 9 are arranged in the activity space, a propeller hub bearing 21 is also arranged between the propeller hub 6 and the fixed ring base 28, the propeller hub bearing 21 is used to reduce the friction force when the propeller hub 6 rotates, a propeller hub bearing 21 cover is arranged on the side of the propeller hub bearing 21 away from the propeller hub 6, and a rotating ring cover plate 22 is arranged on the side of the propeller hub bearing 21 cover away from the propeller hub bearing 21 A bearing 23 and a fixed ring bearing seat 46 for fixing the bearing 23 are also arranged between the rotating ring cover plate 22 and the fixed ring base 28. A bearing 23 is also arranged between the fixed ring bearing seat 46 and the rotating ring cover plate 22. A rotating ring 26 is arranged between the bearing 23 and the fixed ring base 28. A ball head 27 for reducing the rotational friction of the rotating ring 26 is arranged between the rotating ring 26 and the fixed ring base 28. By arranging multiple levels of bearings 23 and rotating rings 26 in the fixed ring base 28, the rotational friction of the hub 6 during rotation is reduced.

[0049] A steering gear mounting seat clamp 30 is connected to the side of the fixing ring base 28 facing away from the propeller hub 6, and a steering gear upper mounting seat 34 is connected to the side of the steering gear mounting seat clamp 30 facing away from the fixing ring base 28. An installation interval for installing the steering gear 32 is formed between the steering gear mounting seat clamp 30 and the steering gear upper mounting seat 34. A guide limit plate 33 is also provided between the steering gear mounting seat clamp 30 and the steering gear upper mounting seat 34. The steering gear 32 is connected to the steering gear upper mounting seat 34. The steering gear upper mounting seat 34 is also provided with a steering gear mounting seat clamp 30 for fixing the steering gear 32, so that the steering gear 32 can be stably installed between the steering gear mounting seat clamp 30 and the steering gear upper mounting seat 34. A steering gear rocker arm 31 is also installed on the steering gear 32, and the steering gear rocker arm 31 is connected to the propeller hub 6, so that the steering gear 32 drives the steering gear rocker arm 31 to drive the propeller hub 6 for adjustment.

[0050] Working principle and its effect:

[0051] The multifunctional test bench for electric rotor configuration characteristics is mainly composed of a column, a test bench, a rotor 13 system, a yaw control mechanism and a pitch control mechanism. The multifunctional test bench can measure the aerodynamic characteristics of a single rotor 13 and multiple rotors 13 under different numbers of rotors 13, different front-to-back spacing, different left-to-right spacing, different height spacing, and different flight attitudes. The rotor 13 system can realize the adjustment of cyclic pitch and total pitch, and the tilt mechanism realizes the horizontal and vertical state switching of the rotor 13 system. The entire tilt rotor 13 system can realize the transformation of spatial position along the crossbeam, longitudinal beam and vertical beam. The yaw angle adjustment is realized by the worm gear mechanism 65, and the pitch direction angle adjustment of the test bench is realized by the electric push rod.

[0052] The rotor 13 system is mainly composed of a pitch mechanism and a tilt mechanism, which is installed on the outside and connected to the rotor 13 installed on the top through a series of push-pull rods. The pitch rocker arms of the blades are connected by a fixed-length pull rod. Therefore, the blades are cyclically pitched by changing the controlled tilt, and the total pitch is adjusted by moving upward or downward. The tilt of the rotor 13 system is mainly changed by the rotation of the tilt servo 44, which causes the servo rocker arm 2 51 to change accordingly. The servo rocker arm 31 is rotated accordingly through the servo rocker arm 1 50 and the servo rocker arm 2 51, and the tilt shaft 49 rotates with the rotation of the servo rocker arm 31, thereby driving the tilt of the rotor 13 system.

[0053] The pitch control system mainly changes the pitch angle of the entire platform by extending and retracting the electric push rod 64 to simulate the static or dynamic aerodynamic characteristics of the aircraft in the pitch state. The worm gear structure is rotated by the planetary reducer 66 and the servo motor 67 to drive the upper platform to rotate to simulate the static or dynamic aerodynamic characteristics of the aircraft in the yaw state.

[0054] The multifunctional test bench adopts a modular design, which is easy to disassemble, assemble and debug. It can measure the aerodynamic characteristics of different numbers and different spatial positions of rotors 13 according to the test requirements. Each rotor 13 system is connected to the test bench by screws, which can be easily disassembled and installed. At the same time, in addition to the pitch and tilt functions, the lateral spacing, longitudinal spacing and height position between each rotor 13 system can be adjusted according to the test requirements.

[0055] The base 68 of the motor 5 is connected to the six-component balance 3 through a balance connection, and the aerodynamic force, anti-torque and tilting moment of the rotor 13 in each direction are measured by the six-component balance 3. Parameters such as rotation speed, current and voltage are measured by sensors, and the cyclic pitch angle, total pitch angle, tilt angle, yaw angle and pitch angle are obtained by calibration calculation of the steering gear 32.

[0056] Embodiment 2:

[0057] The difference from the first embodiment is that the multifunctional test bench system for testing the characteristics of the electric rotor configuration is a specific implementation method of the present invention. The multifunctional test bench system includes a flight dynamics module, a sensor module, a simulation model integration module, a test instance module and a data analysis module, and also includes a flight controller to be tested. Except for the data analysis module, the other four modules are all running in a real-time embedded single-chip microcomputer. Among them, the test instance module generates an excitation signal and sends it to the flight controller to be tested; the flight controller sends a control signal to the flight dynamics module through the simulation model integration module; the flight dynamics module generates angular motion and linear motion information of the multi-rotor aircraft and sends it to the sensor module; the sensor module sends the angular motion and linear motion information of the aircraft with noise to the flight controller; the flight controller sends the expected signal and feedback signal to the data analysis module to form a complete closed-loop data flow.

[0058] The present invention constructs the above test system to carry out flight control testing and prototype manufacturing concurrently, and performs simulation testing on the flight control system in a simulation environment to verify / correct the flight control logic. The complete flight control system logic test includes the test of the attitude loop and the position loop, which can significantly save the research and development cycle.

[0059] The entire test system construction process is the following steps S1-S3, and the test implementation process is the following steps S4-S5.

[0060] S1. Establish the flight dynamics model of multi-rotor aircraft:

[0061] The three-dimensional digital model of the multi-rotor aircraft is used to determine the input parameters of its flight dynamics model, and calculate the dynamics module and kinematics module of its dynamics model.

[0062] Specifically, the input parameters include: total mass of the aircraft m, X-axis moment of inertia J xx, Y-axis moment of inertia J yy 、Z-axis moment of inertia J zz 、XY axis inertia product J xy 、XZ axis inertia product J xz 、YZ axis inertia product J yz , the distance L between the rotor center and the fuselage center on the transverse channel roll , the distance L between the rotor center and the fuselage center on the transverse channel pitch , the maximum lift generated by a single rotor T max , the maximum torque Q generated by a single rotor max 、Fuselage drag coefficient C f 、Fuselage damping moment coefficient C m .

[0063] The input signal of the multi-rotor aircraft dynamics model received from the flight control system is the PWM wave signal of each motor. The calculation steps from each motor PWM signal to the total thrust and torque of the multi-rotor in the n-type rotor are:

[0064] T B =T max ·(u1+u2+...+u n )

[0065]

[0066] Among them: Input parameter u i (i=1...n) is the normalized value of each motor PWM, n represents the number of motors in the multi-rotor aircraft (the n value for the X-type six-rotor aircraft is 6), α i (i=1...n) is the roll coefficient of each rotor of the multi-rotor aircraft to generate a roll moment, β i (i=1...n) is the pitch coefficient of each rotor of the multi-rotor aircraft to generate the pitch moment, γ i (i=1...n) The yaw coefficient of each rotor of the multi-rotor aircraft to generate the yaw moment, output variable T B , They are the total tension and the moments around the three principal axes in the body coordinate system respectively.

[0067] Next, the dynamics module of the multi-rotor aircraft dynamics model is calculated, and the calculation steps are as follows:

[0068]

[0069] Among them, m represents the weight of the multi-rotor aircraft, u, v, and w represent the three linear velocities of the multi-rotor aircraft in the body coordinate system, and p, q, and r represent the three angular velocities of the multi-rotor aircraft around the body coordinate system. and Denote the derivatives of .

[0070] Then, the kinematic module of the multi-rotor aircraft dynamics model is calculated, and the calculation steps are as follows:

[0071]

[0072] in, is the three-dimensional position coordinate in the inertial coordinate system, φ, θ, ψ are the Euler angles of the aircraft, Represents the transformation matrix from the body coordinate system to the inertial coordinate system.

[0073] S2. Build digital models of sensors including accelerometer, gyroscope, magnetometer, barometer and GPS.

[0074] First, the calculation model of the three-axis accelerometer is established. The calculation steps are:

[0075]

[0076] Among them, sensor acc is the output of the accelerometer model, η acc is the Gaussian white noise of the accelerometer, g is the gravity constant, Represents the linear acceleration in the body coordinate system, Ω B Represents the angular velocity vector in the body coordinate system, V B represents the linear velocity in the body coordinate system, Represents the transformation matrix from the inertial coordinate system to the body coordinate system.

[0077] Next, the calculation model of the three-axis gyroscope is established. The calculation steps are as follows: sensor gyr =Ω B +η gyr ,

[0078] Among them, sensor gyr is the output of the gyroscope model, η gyr is the Gaussian white noise of the gyroscope, Ω B is the angular velocity vector in the body coordinate system.

[0079] Then, the calculation model of the three-axis magnetometer is established, and its calculation steps are: Among them, sensor mag is the output of the magnetometer model, η mag is the Gaussian white noise of the magnetometer, It is the transformation matrix from the inertial coordinate system to the body coordinate system.

[0080] Next, the calculation model of the barometer is established. The calculation steps are as follows: sensor bar =pe +η bar , where sensor bar is the output of the barometer model, η bar is the Gaussian white noise of the barometer, p e is the actual height of the multirotor aircraft.

[0081] Finally, the calculation model of GPS is established, and its calculation steps are as follows: sensor gps =f (x,y,z) +η gps , where sensor gps Here is the output of the GPS model, f (x,y,z) is the Gaussian white noise of GPS, η gps is the actual position of the multirotor aircraft.

[0082] S3. Deploy the established multi-rotor digital model to the flight controller platform to be tested. FCU =Sign SENSOR Integration process for simulation models.

[0083] The specific implementation process is:

[0084] First, perform system-level simulation on a desktop computer to verify the performance parameters of the flight dynamics model of the established multi-rotor model and record the input and output data. It should be noted that the test results of the model-in-the-loop simulation must be correct before entering the next step of running the established flight dynamics model and sensor model in a real-time computer environment. Specifically, convert the models obtained in steps S1 and S2 into C / C++ code, deploy the model code to a real-time microcontroller, and then test the controller as part of the actual aircraft control system.

[0085] Among them, the interface between the flight controller and the flight dynamics model is: PWM i (i=1...n), the interface between the sensor model and the flight controller is:

[0086] FCU acc =sensor acc

[0087] FCU gyr =sensor gyr

[0088] FCU mag =sensor mag

[0089] FCU bar =sensor bar

[0090] FCUgps =sensor gps

[0091] Among them, FCU stands for flight controller, the subscript acc stands for accelerometer, gyr stands for gyroscope, mag stands for magnetometer, bar stands for barometer, and gps stands for GPS positioning module; sensor stands for sensor, and its meaning in the table below is similar to the above.

[0092] FCU acc =sensor acc It means transmitting the data of the acceleration sensor to the accelerometer calculation module of the flight controller, and the others are similar.

[0093] In this way, the flight controller, flight dynamics model, and sensor model form a closed loop, and all three run in a real-time computer, making the flight controller closer to the real environment and the simulation results more accurate and reliable.

[0094] S4. Input the test case and compare the simulation test results with the results output by the flight controller hardware platform to be tested. First, test whether the attitude control mode control logic is correct.

[0095] The initial conditions are set as:

[0096] Among them, h0 indicates that the initial height of the multi-rotor aircraft is 10 meters, θ0 indicates that the initial pitch angle is 0, φ0 indicates that the initial roll angle is 0, ψ0 indicates that the initial yaw angle is 0, and PWM i (i=1...n)=1500 means that the initial control signal PWM of the first motor is 1500.

[0097] The steps to test the pitch channel of the attitude control mode are as follows: Set the real-time flight control system input signal to θ d = ±30°, and then record the real-time attitude angle data: φ, θ, ψ in the flight data recording module. The recording time is 10s, and the obtained data is named TestAttPitch.

[0098] The steps to test the roll channel of the attitude control mode are as follows: Set the real-time flight control system input signal to φ d =±30°, and then record the real-time attitude angle data: φ, θ, ψ in the flight data recording module. The recording time is 10s, and the obtained data is named TestAttRoll.

[0099] The steps to test the attitude control mode yaw channel are: set the real-time flight control system input signal to ψ d=±30°, and then record the real-time attitude angle data: φ, θ, ψ in the flight data recording module. The recording time is 10s, and the obtained data is named TestAttYaw.

[0100] Then, test whether the position control mode control logic is correct.

[0101] The initial conditions are set as:

[0102] Among them, h0=10 means that the initial height of the multi-rotor aircraft is 10 meters, p n0 =0, p e0 =0 means the initial horizontal position is 0, θ0=0 means the initial pitch angle is 0, φ0=0 means the initial roll angle is 0, ψ0=0 means the initial yaw angle is 0, PWM i (i=1...n)=1500 means that the initial control signal PWM of the first motor is 1500.

[0103] The steps for testing the horizontal and longitudinal channels of the position control mode are as follows: Set the real-time flight control system input signal to p nd = ±1, indicating that the expected longitudinal horizontal displacement varies within plus or minus one meter, and then the real-time position and attitude angle data are recorded in the flight data recording module: p n 、p e , h, φ, θ, ψ, the recording time is 120s, and the obtained data is named TestPosNorth.

[0104] The steps to test the horizontal lateral channel of the position control mode are as follows: Set the real-time flight control system input signal to p ed = ±1, indicating that the expected lateral horizontal displacement varies within plus or minus one meter, and then the real-time position and attitude angle data are recorded in the flight data recording module: p n 、p e , h, φ, θ, ψ, the recording time is 120s, and the obtained data is named TestPosEast.

[0105] The steps to test the altitude channel of the position control mode are as follows: Set the real-time flight control system input signal to h d =35, indicating that the expected altitude is 35m, and then the real-time position and attitude angle data are recorded in the flight data recording module: p n 、p e , h, φ, θ, ψ, the recording time is 15s, and the obtained data is named TestPosHeight.

[0106] S5: This step is Figure 1The flight control logic judgment module in the implementation method is to analyze the test database obtained in step S4, which is specifically composed of the TestAttPitch, TestAttRoll, TestAttYaw, TestPosNorth, TestPosEast, and TestPosHeight data packets obtained in the previous step S4.

[0107] The essence of this step is to modify the control logic of the multi-rotor aircraft control system according to the test results, and then return to step S4 to iterate repeatedly. d and the expected signal Sign of the actual feedback real If the change pattern of is consistent, it proves that the control logic of the channel is correct. If the change pattern is completely opposite, it is necessary to return to step S4 to correct the flight control logic of the flight controller. In order to better understand the present invention, the roll channel of the attitude control mode is used as an example for explanation: First, compare the expected roll angle signal φ d If the actual feedback roll angle signal φ has the same change pattern as the expected roll angle signal within 10s, it proves that the control logic of the roll channel in the attitude mode is correct; if the change pattern is completely opposite, it is necessary to return to step S4 to correct the flight control logic of the flight controller. The flight control logic tests of other channels and each channel in other modes are similar and will not be repeated here.

[0108] A schematic diagram of the test results of the horizontal position controller of an aircraft control system provided for this embodiment, in which the solid black line represents the expected horizontal position signal, and the dotted black line represents the actual multi-rotor horizontal position. It can be seen that the changing trends of the two are generally consistent, indicating that the logic design of the tested horizontal position controller is correct; if the changing trends of the two are very different or completely opposite, it means that the logic design of the tested controller is wrong, and the aircraft flight control system designer needs to correct it and retest it after correction until the changing trends of the two are basically consistent, thus terminating the test instance.

[0109] It can be seen that the testing method of the present invention can adapt to the flight control system testing requirements of the aircraft, improve the performance of the aircraft, and ensure the safety of the aircraft in subsequent actual flight processes. In addition, the testing method can be carried out concurrently with the manufacturing department, which can greatly shorten the R&D cycle and reduce R&D costs.

[0110] The above is only a preferred embodiment of the present invention, and the protection scope of the present invention is not limited to the above embodiments. All technical solutions under the concept of the present invention belong to the protection scope of the present invention. It should be pointed out that for ordinary technicians in this technical field, some improvements and modifications without departing from the principle of the present invention should also be regarded as the protection scope of the present invention.

Claims

1. A multifunctional test bench for electric rotor configuration characteristics, characterized by: Including the bench base, bench platform and tilting mechanism, The gantry platform is rotatably connected to the gantry base, the tilt mechanism is arranged on the gantry platform, the tilt mechanism is further connected to a tilt rotor (13), and the gantry base is used to control the rotation and pitch adjustment of the gantry platform; The tilt mechanism is used to control the simulated tilting of the tilt rotor (13) to adjust the angle of the tilt rotor (13), and to simulate the aerodynamic characteristics of the aircraft in a pitch state, static or dynamic flight, and the attitude of the aircraft during flight.

2. The multifunctional test bench for electric rotor configuration characteristics according to claim 1, characterized in that: The gantry base includes a base (68), a pillar (62) and an electric push rod (64); a rotating mechanism is connected between the base (68) and the pillar (62); the rotating mechanism includes a servo motor (67), a planetary reducer (66) and a worm gear mechanism (65); a seat body is provided on the base (68); a worm gear is provided in the seat body; the worm gear is connected to the pillar (62); the planetary reducer (66) is provided on the seat body; the input shaft of the planetary reducer (66) is coaxially fixedly connected to the servo motor (67); the output shaft of the planetary reducer (66) is coaxially fixedly connected to the worm; the worm is meshed with the worm gear.

3. The multifunctional test bench for electric rotor configuration characteristics according to claim 2, characterized in that: The tilt mechanism comprises a fixed beam (36), a reinforcement pipe (38), a C-shaped beam (39), an electric adjustment mounting seat, a tilt servo (44) and a servo rocker arm (31); the fixed beam (36) is connected to the fixed pipe of the rotor (13); the reinforcement pipe (38) is arranged on the fixed beam (36); a reinforcing rib (37) is further connected between the reinforcement pipe (38) and the fixed beam (36); the end of the reinforcement pipe (38) away from the reinforcement pipe (38) is connected to the electric adjustment mounting seat; the electric adjustment mounting seat comprises an electric adjustment mounting seat 1 (40) and an electric adjustment mounting seat 2 (41); the electric adjustment mounting seat 1 (40) and the electric adjustment mounting seat 2 (41) are used to install the electric adjustment (42); the C-shaped beam (39) is located on the side of the reinforcement pipe (38) away from the reinforcing rib (37) and is used to connect the reinforcement pipe (38) and the fixed beam (36).

4. The multifunctional test bench for electric rotor configuration characteristics according to claim 3 is characterized by: The C-shaped beam (39) is further provided with a tilting servo (44) mounting seat, the tilting servo (44) is mounted on the tilting servo (44) mounting seat, the tilting servo (44) fastening seat is used to connect with the tilting motor (5) mounting seat and to fix the tilting servo (44) on the tilting servo (44) mounting seat, the tilting servo (44) is further provided with a bearing seat (46), the bearing seat (46) is provided with a bearing (48) and a clamp (47) for fixing the bearing (48).

5. The multifunctional test bench for electric rotor configuration characteristics according to claim 4 is characterized by: A tilt shaft (49) is provided at the axis position of the bearing (48), and the tilt servo (44) is used to drive the tilt shaft (49) to rotate. The shaft end of the tilt shaft (49) is connected to the servo rocker arm (31), and the servo rocker arm (31) includes a servo rocker arm 1 (50) and a servo rocker arm 2 (51). The servo rocker arm 1 (50) is connected to the tilt shaft (49), and the servo rocker arm 1 (50) and the servo rocker arm 2 (51) are respectively connected to a ball head pull rod (52), and the ball head pull rod (52) between the servo rocker arm 1 (50) and the servo rocker arm 2 (51) are respectively connected to a servo pull rod (53).

6. The multifunctional test bench for electric rotor configuration characteristics according to claim 5, characterized in that: The tilt rotor (13) comprises a tilt connection plate (2), a motor connection member (4), a motor (5), a propeller hub (6) and a rotor (13), wherein the tilt connection plate (2) is connected to a tilt shaft (49), a steel shaft (1) for connecting the motor (5) is further provided in the tilt connection plate (2), a six-component balance (3) is further connected to the motor (5), the steel shaft (1) is connected to the motor connection member (4), the motor (5) is reinforced with the connection to the tilt connection plate (2) through the motor connection member (4), an output end of the motor (5) is connected to the propeller hub (6), a propeller hub bearing (8) is provided between the propeller hub (6) and the rotor (13), the propeller hub bearing (8) is connected to the propeller hub (6), and an adjustment ring (7) is further provided between the propeller hub bearing (8) and the propeller hub (6).

7. The multifunctional test bench for electric rotor configuration characteristics according to claim 6, characterized in that: The propeller hub bearing (8) is connected to a propeller clamp (9) on the side facing away from the propeller hub (6), a thrust bearing (11) is provided between the propeller clamp (9) and the rotor (13), a gasket (12) is provided between the thrust bearing (11) and the rotor (13), an adjustment plate (10) is provided between the thrust bearing (11) and the propeller clamp (9), a tie rod ball buckle (19) is provided between the propeller hub (6) and the propeller clamp (9), the tie rod ball buckle (19) is connected to the propeller hub (6) and the propeller clamp (9), a pitch rod (18) is connected to the tie rod ball buckle (19), and the pitch rod (18) is provided on the pitch rod (18). 8) A pitch rod hinge (17) is connected to one end away from the rod ball buckle (19), and the pitch rod hinge (17) is connected to the bottom of the hub (6) on the side away from the pitch rod (18). A pitch rod hinge bearing (15) is provided at the connection between the pitch rod hinge (17) and the hub (6), a retaining ring (14) is provided between the pitch rod hinge (17) connecting bearing (48) and the hub (6), and a shaft sleeve (16) is also provided between the pitch rod hinge bearing (15) and the retaining ring (14). The hub (6) adjusts the tilt angle of the rotor (13) under the action of the pitch rod (18).

8. The multifunctional test bench for electric rotor configuration characteristics according to claim 7 is characterized by: A steering gear (32) for driving a pitch-varying pull rod (18) is provided on the side of the propeller hub (6) facing away from the motor (5); a fixed ring base (28) is also provided on the motor (5); an activity space is formed between the fixed ring base (28) and the motor (5); the propeller hub (6) and the propeller clamp (9) are arranged in the activity space; a propeller hub bearing (8) is also provided between the propeller hub (6) and the fixed ring base (28); the propeller hub bearing (8) is used to reduce friction when the propeller hub (6) rotates; a propeller hub bearing cover (21) is provided on the side of the propeller hub bearing (8) facing away from the propeller hub (6); a rotating ring cover (22) is provided on the side of the propeller hub bearing cover (21) facing away from the propeller hub bearing (8); a bearing (48) is also provided between the rotating ring cover (22) and the fixed ring base (28), and a fixed ring bearing seat (46) for fixing the bearing (48).

9. The multifunctional test bench for electric rotor configuration characteristics according to claim 8, characterized in that: A bearing (48) is further provided between the fixed ring bearing seat (46) and the rotating ring cover plate (22), a rotating ring (26) is provided between the bearing (48) and the fixed ring base (28), and a ball head (27) for reducing the rotational friction of the rotating ring (26) is provided between the rotating ring (26) and the fixed ring base (28).

10. A test system for a multifunctional test bench for electric rotor configuration characteristics, applied to the multifunctional test bench for electric rotor configuration characteristics as claimed in any one of claims 1 to 9, characterized in that: The system comprises a flight dynamics module, a sensor module, a simulation model integration module, a test case module and a data analysis module. The five modules constitute a closed-loop data flow, and the flight dynamics module, the sensor module, the simulation model integration module and the test case module all run on a real-time embedded single-chip microcomputer. The test instance module generates an excitation signal and sends it to the flight controller to be tested; The simulation model integration module sends the control signal of the flight controller to be tested to the flight dynamics module; The flight dynamics module generates angular motion and linear motion information of the multirotor aircraft and sends it to the sensor module; The sensor module sends the aircraft angular motion and linear motion information with noise to the flight controller; The data analysis module receives the expected signal and the feedback signal fed back by the flight controller to be tested.

Images