Rotor wing tilting mechanism of electric vertical take-off and landing aircraft and working method

Abstract

The invention belongs to the technical field of aviation, and particularly relates to a rotor wing tilting mechanism of an electric vertical take-off and landing aircraft and a working method. According to the tilting mechanism, the whole tilting mechanism is arranged in a bilateral symmetry mode, in addition, the left side component and the right side component are connected through the concentric shaft which is transversely arranged, the strength and rigidity of the whole tilting mechanism in the movement process are enhanced, vibration in the complex flight state can be effectively reduced through stable mechanical arrangement, and the tilting precision is guaranteed. A unique connecting rod mechanism is adopted, so that the tilting rotor wings in the middle sections of the wings only tilt blades, a cabin body mechanism where the rotor wings are located does not need to integrally tilt, and a power redundancy design driven by double linear motors is adopted; the tilting mechanism provides a new tilting mode selection for the tilting type electric vertical take-off and landing aircraft, avoids the problems of wing additional lower washing load caused by overall tilting of the nacelle and unstable large stall attack angle of the wing tilting mode, and is beneficial for the aircraft to give full play to the advantages of vertical take-off and landing, high-speed forward flight and the like.

Description

Technical Field

[0001] The invention belongs to the field of aviation technology, and in particular relates to a rotor tilting mechanism and a working method of an electric vertical take-off and landing aircraft. Background Art

[0002] With the rapid development of battery technology and electric propulsion systems, the endurance and flight performance of electric aircraft are constantly improving, providing a new direction for future air transportation. Electric aircraft are powered by electricity, and compared with traditional fuel aircraft, their greenhouse gas emissions are significantly reduced, which can effectively reduce air pollution and noise pollution, help address climate change and improve the urban environment. More and more countries and regions have set goals to reduce carbon emissions, and the market demand for green aviation solutions has increased, which has promoted the research and development and application of electric vertical aircraft.

[0003] At present, electric vertical take-off and landing aircraft mainly have multi-rotor, composite wing, tilt-rotor and other configurations. The rotors of multi-rotor aircraft are highly efficient during vertical take-off and landing and hovering, but are less efficient during forward flight and consume more energy; due to energy efficiency issues, the endurance is usually poor. The lift rotors of composite wing aircraft will generate greater resistance during cruising, reducing the flight speed and range of the entire aircraft; the full power unit is not effectively utilized, which in turn increases the weight of the entire aircraft structure. Tilt-rotor aircraft combines the advantages of rotorcraft and fixed-wing aircraft, and have the advantages of vertical take-off and landing, hovering in the air, fast forward flight speed and long range.

[0004] The tilt mechanism of a tilt-type electric vertical take-off and landing aircraft usually adopts a slider screw and gear transmission form. Its overall structure is complex, which is not conducive to later maintenance and pre-flight and post-flight inspections. The above transmission accuracy will decrease after repeated use in both directions, and it is difficult to meet the high-precision control requirements of the tilt mechanism. The power redundancy safety design of the tilt mechanism is not considered. When a single tilt servo fails, the tilt system will become uncontrollable and flight safety cannot be guaranteed. In the fixed-wing aircraft mode, the aerodynamic shape of the rotor fairing and the fuselage skin is discontinuous, which is not conducive to high-speed forward flight performance.

[0005] Therefore, in order to solve the problems mentioned in the background technology, such as insufficient consideration of safety redundancy design of the tilt mechanism, difficulty in ensuring the transmission accuracy of the screw and gear, and discontinuity of the aerodynamic shape, the present invention provides an electric vertical take-off and landing aircraft rotor tilt mechanism and a working method. Summary of the invention

[0006] The objects of the present invention are:

[0007] The purpose of the present invention is to provide an electric vertical take-off and landing aircraft rotor tilting mechanism and working method to solve the problems mentioned in the background technology, such as insufficient consideration of safety redundancy design, difficulty in ensuring the transmission accuracy of the screw and gear, and discontinuity of the aerodynamic shape.

[0008] To achieve the above-mentioned purpose, the present invention provides the following technical solutions: an electric vertical take-off and landing aircraft rotor tilting mechanism, comprising a body system 100, a drive system 200, a transmission system 300, and a rotor system 400; the drive system 200 comprises a linear motor a201, a linear motor b202, a cam a203, and a cam b204; the front ends of the linear motors a201 and b202 are respectively connected to the ears of the cams a203 and b204 through bearings, and the end mounting points of the linear motors a201 and b202 are both hinged to the rotor arm 101 of the body system 100, and the cams a203 and b204 are respectively connected to the splints a102 and b103 through bearings; the transmission system 300 comprises a cam connecting rod assembly 301, a diagonal rod assembly 302, a diamond rod assembly 303, a curved rod 304, a support plate 305, a cam connecting rod assembly 306, a cam connecting rod assembly 307, a cam connecting rod assembly 308, a cam connecting rod assembly 309, a cam connecting rod assembly 310, a cam connecting rod assembly 311, a cam connecting rod assembly 312, a cam connecting rod assembly 313, a cam connecting rod assembly 314, a cam connecting rod assembly 315, a cam connecting rod assembly 316, a cam connecting rod assembly 317, a cam connecting rod assembly 318, a cam connecting rod assembly 319, a cam connecting rod assembly 320 The rod assembly concentric shaft 306, the inclined rod assembly concentric shaft 307, the middle concentric shaft 308, and the crank rod concentric shaft 309; one end of the cam connecting rod assembly 301 is hinged to the ends of the cam a203 and the cam b204 through the cam connecting rod assembly concentric shaft 306, one end of the inclined rod assembly 302 is connected to the clamping plate a102 and the clamping plate b103 of the body system 100 through the inclined rod assembly concentric shaft 307, and one end of the diamond rod assembly 303 is connected to the middle The coaxial shaft 308 is connected to the cam-link assembly 301 and the inclined rod assembly 302 by bearings, the diamond rod assembly 303 is hinged to the bent rod 304 at its middle part by a bearing, the left end of the bent rod 304 is connected to the clamping plate a102 and the clamping plate b103 by the bent rod coaxial shaft 309, and the support plate 305 is hinged to the right end of the bent rod 304; the motor support 403 of the rotor system 400 is hinged to the diamond rod assembly 303 and the support plate 305 by bearings.

[0009] Furthermore, the body system 100 includes a rotor arm 101, a splint a102, a splint b103, and a rotor arm skin 104; the splint a102 and the splint b103 are fixedly connected to the right end of the rotor arm 101, and are arranged symmetrically on the left and right, and the rotor arm skin 104 covers the outside of the rotor arm 101.

[0010] Furthermore, the rotor arm 101 is an I-shaped beam box structure, the right end face of the rotor arm 101 is provided with a helicopter mode limiting slope 1011 and a fixed-wing aircraft mode limiting slope 1012 of the tilt mechanism, and the left end of the rotor arm 101 is fixedly connected to the wing load-bearing structure.

[0011] Furthermore, the splint a102 includes a self-locking pin a1021 and an electromagnetic spring a1022; the self-locking pin a1021 and the electromagnetic spring a1022 are fixedly connected; the splint b103 contains a self-locking pin b1031 and an electromagnetic spring b1032, the self-locking pin b1031 and the electromagnetic spring b1032 are fixedly connected, and the self-locking pin a1021 and the self-locking pin b1031 are respectively arranged on the splint a102 and the splint b103 in a distributed circle.

[0012] Furthermore, the linear motor a201 and the linear motor b202 have built-in dual Hall elements.

[0013] Furthermore, the sizes of the cam a203 and the cam b204 and the positions of their ears always meet the minimum force arm value requirements of the linear motor a201 and the linear motor b202 during the tilting process.

[0014] Furthermore, the linear motor a201 includes a motor base a2011 and a retractable push rod a2012; the retractable push rod a2012 is installed inside the motor base a2011, and the linear motor b202 includes a motor base b2021 and a retractable push rod b2022; the retractable push rod b2022 is installed inside the motor base b2021.

[0015] Further, the cam a203 has a helicopter mode limiting slope a2031 and a fixed-wing aircraft mode limiting slope a2032, and the cam b204 has a helicopter mode limiting slope b2041 and a fixed-wing aircraft mode limiting slope b2042. In the helicopter mode, the helicopter mode limiting slope a2031 and the helicopter mode limiting slope b2041 are in contact with the helicopter mode limiting slope 1011 of the rotor arm 101; in the fixed-wing aircraft mode, the fixed-wing aircraft mode limiting slope a2032 and the fixed-wing aircraft mode limiting slope b2042 are in contact with the fixed-wing aircraft mode limiting slope 1012 of the rotor arm 101.

[0016] Furthermore, a circle of self-locking holes a2033 are distributed on the surface of cam a203, and a circle of self-locking holes b2043 are distributed on the surface of cam b204. In the event of an accidental power outage, in addition to the linear motor being self-locking due to power failure, the self-locking pin a1021 will be inserted into the self-locking hole a2033 under the action of the electromagnetic spring a1022; the self-locking pin b1031 will be inserted into the self-locking hole b2043 under the action of the electromagnetic spring b1032, and together with the linear motor, prevent the tilting mechanism from rotating at will.

[0017] Furthermore, the cam-link assembly 301 includes a cam link a3011 and a cam link b3012 which are symmetrically arranged on the left and right; one end of the cam link a3011 and the cam link b3012 are respectively hinged to the cam a203 and the cam b204 through the coaxial shaft 306 of the cam-link assembly, the other end of the cam link a3011 is hinged to the diagonal rod a3021 and the diamond rod a3031 through the middle concentric shaft 308, and the other end of the cam link b3012 is hinged to the diagonal rod b3022 and the diamond rod b3032 through the middle concentric shaft 308.

[0018] Furthermore, the diagonal tie rod assembly 302 includes a diagonal tie rod a3021 and a diagonal tie rod b3022 which are symmetrically arranged on the left and right; one end of the diagonal tie rod a3021 and the diagonal tie rod b3022 are connected to the coaxial axis 307 of the diagonal tie rod assembly by a bearing, and the other end is hinged to the cam link a3011, the diamond rod a3031, the cam link b3012 and the diamond rod b3032 through the middle coaxial axis 308, and the coaxial axis 307 of the diagonal tie rod assembly is connected to the splint a102 and the splint b103 for anti-rotation.

[0019] Furthermore, the diamond rod assembly 303 includes a diamond rod a3031 and a diamond rod b3032 that are symmetrically arranged on the left and right; the left end of the diamond rod a3031 is hinged to the cam connecting rod a3011 and the diagonal rod a3021 through the middle concentric shaft 308, and the left end of the diamond rod b3032 is hinged to the cam connecting rod b3012 and the diagonal rod b3022 through the middle concentric shaft 308, and the diamond rod a3031 and the diamond rod b3032 are connected to both sides of the curved rod 304 through bearings in the middle.

[0020] Furthermore, the left end of the curved rod 304 is connected to the curved rod concentric axis 309 through a bearing, and the right end is hinged to the support plate 305. The curved rod concentric axis 309 and the clamping plates a102 and b103 are anti-rotationally connected.

[0021] Furthermore, rotor system 400 includes fairing 401, rotor drive motor 402, motor support 403, rotor hub 404, and rotor blades 405; rotor drive motor 402 and motor support 403 are installed inside fairing 401, rotor drive motor 402 is connected to rotor hub 404, and rotor blades 405 are installed on rotor hub 404.

[0022] Furthermore, the lower end surface of the motor support 403 has four groups of ears, including ear a4031, ear b4032, ear c4033 and ear d4034. Motor heat dissipation holes are distributed on the motor support 403. Ear a4031 and ear b4032 are hinged to the diamond rod b3032 and diamond rod a3031 respectively. Ear c4033 and ear d4034 are hinged to the support plate 305 through bearings.

[0023] A working method of a rotor tilting mechanism of an electric vertical take-off and landing aircraft,

[0024] After receiving the command from the flight control system, the linear motor a201 and the linear motor b202 drive the retractable push rod a2012 and the retractable push rod b2022 to retract, thereby driving the cam a203 and the cam b204 to rotate. The rotation of the cam a203 and the cam b204 causes the synchronous movement and rotation of the transmission system 300. The transmission system 300 gradually transmits the input of the drive system 200 to the output end component of the transmission system 300, and finally realizes the conversion of the rotor system 400 between the helicopter mode and the fixed-wing aircraft mode.

[0025] When the power is on, the electromagnetic spring a1022 and the electromagnetic spring b1032 are in a compressed state due to the magnetic force, pulling the self-locking pin a1021 and the self-locking pin b1031 to shrink within the plane of the clamping plate a102 and the clamping plate b103; when the power is off, the electromagnetic spring a1022 and the electromagnetic spring b1032 lose the magnetic force restriction and are in an extended state, and the self-locking pin a1021 and the self-locking pin b1031 exceed the plane of the clamping plate a102 and the clamping plate b103;

[0026] If the linear motor a201 / linear motor b202 loses power accidentally during the tilting process, the instantaneous power of the linear motor b202 / linear motor a201 is increased to keep the tilting motion going. If the linear motor a201 and the linear motor b202 lose power accidentally at the same time, the self-locking pin a1021 will be inserted into the self-locking hole a2033 under the action of the electromagnetic spring a1022. The self-locking pin b1031 will be inserted into the self-locking hole b2043 under the action of the electromagnetic spring b1032, further enhancing the self-locking effect on the basis of the self-locking function of the linear motor.

[0027] In helicopter mode, the helicopter mode limit slope a2031 and the limit slope b2041 are in contact with the helicopter mode limit slope 1011 of the rotor arm 101; in fixed-wing aircraft mode, the fixed-wing aircraft mode limit slope a2032 and the limit slope b2042 are in contact with the fixed-wing aircraft mode limit slope 1012 of the rotor arm 101. In addition to the position control of the linear motor a201 and the linear motor b202, the above-mentioned method further limits excessive tilting of the tilt mechanism at the two extreme positions.

[0028] The technical effects and advantages of the present invention are as follows: the tilt mechanism provides a new tilt mode option for tilt-type electric vertical take-off and landing aircraft, avoiding the additional downwash load on the wing caused by the overall tilting of the nacelle and the instability problem of large stall angle of attack in the form of wing tilting, which is conducive to this type of aircraft to give full play to the advantages of vertical take-off and landing and high-speed forward flight.

[0029] For electric vertical take-off and landing aircraft with tilt-rotors in the middle of the wing, it is difficult to tilt these rotors without tilting the entire wing. The present invention adopts a unique connecting rod mechanism to allow the tilt-rotors in the middle of the wing to only tilt the blades without tilting the entire cabin mechanism where the rotors are located.

[0030] During the tilting process of the tilting mechanism, the tilting part rotates and moves synchronously instead of rotating around a certain axis, which ensures the continuity of the rotor fairing and the rotor arm skin in the fixed-wing flight mode, giving it a good aerodynamic shape and improving the aerodynamic performance of the entire aircraft.

[0031] The tilt mechanism driver is a high-precision linear motor with power-off self-locking function. It has the advantages of low noise and low vibration, while ensuring flight safety and tilt accuracy. The dual linear motor drive power redundancy design is adopted. When an accident occurs to one motor, the other motor can increase the instantaneous power to ensure that the tilting process continues. The limit system and mechanical power-off self-locking system are set to further enhance the safety redundancy design.

[0032] The main body of the present invention is designed based on a connecting rod mechanism. The rod moves in a plane during the tilting process. All the moving pairs are in surface contact. The force per unit area during transmission is small, and it is conducive to lubrication, so the wear is light and the service life is long. In addition, since the contact surface is mostly a cylindrical surface or a plane, the manufacturing is simpler and it is easy to obtain a higher manufacturing accuracy.

[0033] From the linear motors of the drive system to the structural parts of the transmission system, all selected components or systems have low operating noise, which fully meets the "low noise" concept of electric vertical take-off and landing aircraft.

[0034] The mechanical structure of the tilting mechanism is clear, the positions and functions of each component are clear and easy to observe, and standardized interfaces and connections are adopted, which makes disassembly and maintenance easy. It can not only improve the convenience of operation, but also significantly improve the efficiency of pre-flight and post-flight inspections.

[0035] The tilt mechanism is generally arranged symmetrically on the left and right. In addition, the left and right components are connected by transversely arranged concentric axes, which enhances the strength and rigidity of the entire tilt mechanism during movement. The stable mechanical layout can also effectively reduce vibrations under complex flight conditions and ensure tilt accuracy. BRIEF DESCRIPTION OF THE DRAWINGS

[0036] Figure 1 Schematic diagram of the present invention; wherein: a is a helicopter mode, b is a fixed-wing aircraft mode;

[0037] Among them: 100-airframe system, 200-drive system, 300-transmission system, 400-rotor system.

[0038] Figure 2It is a top view of the present invention; wherein: a is a helicopter mode, and b is a fixed-wing aircraft mode.

[0039] Figure 3 It is a front view of the present invention; wherein: a is a helicopter mode, and b is a fixed-wing aircraft mode.

[0040] Figure 4 It is a side view of the present invention; wherein: a is a helicopter mode, and b is a fixed-wing aircraft mode.

[0041] Figure 5 Schematic diagram of the airframe system 100 of the present invention; wherein: a is a stereoscopic diagram of the airframe system 100, b is a schematic diagram of the rotor arm skin, and c is a front view of the airframe system 100;

[0042] Wherein: 101-rotor arm, 102-clamp a, 103-clamp b, 104-rotor arm skin, 1011-helicopter mode limiting slope, 1012-fixed-wing aircraft mode limiting slope.

[0043] Figure 6 Schematic diagram of the drive system 200 of the present invention; wherein: a is a stereoscopic diagram of the drive system 200, and b is a schematic diagram of the structure of 201-linear motor a, 202-linear motor b, 203-cam a, and 204-cam b;

[0044] Wherein: 201-linear motor a, 202-linear motor b, 203-cam a, 204-cam b, 2011-motor base a, 2021-motor base b, 2012-retractable push rod a, 2022-retractable push rod b, 2031-helicopter mode limiting slope a, 2041-helicopter mode limiting slope b, 2032 fixed-wing aircraft mode limiting slope a, 2042-fixed-wing aircraft mode limiting slope b, 2033-self-locking hole a, 2043-self-locking hole b.

[0045] Figure 7 Schematic diagram of the transmission system 300 of the present invention; wherein: a is a stereoscopic diagram of the transmission system 300, and b is a schematic diagram of the cam connecting rod assembly 301, the inclined rod assembly 302, and the diamond rod assembly 303;

[0046] Among them: 301-cam connecting rod assembly, 302-oblique tie rod assembly, 303-diamond rod assembly, 304-bend rod, 305-support plate, 306-cam connecting rod assembly concentric axis, 307-oblique tie rod assembly concentric axis, 308-middle concentric axis, 309-bend rod concentric axis, 3011-cam connecting rod a, 3012-cam connecting rod b, 3021-oblique tie rod a, 3022-oblique tie rod b, 3031-diamond rod a, 3032-diamond rod b.

[0047] Figure 8It is a schematic diagram of a rotor system 400 of the present invention; wherein: a is a schematic diagram of a fairing 401, b is a schematic diagram of a rotor drive motor 402, a motor support 403, a rotor hub 404, and a blade 405, and c is a schematic diagram of a motor support ear.

[0048] Wherein: 401- fairing, 402- rotor drive motor, 403- motor support, 404- rotor hub, 405- blade, 4031- ear piece a, 4032- ear piece b, 4033- ear piece c, 4034- ear piece d.

[0049] Fig. 9 It is a schematic diagram of the continuity of the aerodynamic shape of the present invention; wherein: a is the helicopter mode, and b is the fixed-wing aircraft mode.

[0050] Fig.10 It is a schematic diagram of the symmetry and power redundancy of the tilting mechanism of the present invention.

[0051] Fig.11 The diagram is a schematic diagram of the limit position of the present invention; wherein: a is a helicopter mode, and b is a fixed-wing aircraft mode.

[0052] Fig.12 Schematic diagram of the splint-cam mechanical self-locking; where a is the power-on state and b is the power-off state.

[0053] Among them: 1021-self-locking pin a, 1031-self-locking pin b, 1022-electromagnetic spring a, 1032-electromagnetic spring b. DETAILED DESCRIPTION

[0054] The present invention is further described below in conjunction with the accompanying drawings:

[0055] like Figure 1-Figure 7As shown, an electric vertical take-off and landing aircraft rotor tilting mechanism includes a body system 100, a drive system 200, a transmission system 300, and a rotor system 400; the drive system 200 includes a linear motor a201, a linear motor b202, a cam a203, and a cam b204; the front ends of the linear motors a201 and b202 are respectively connected to the ears of the cams a203 and b204 through bearings, and the end mounting points of the linear motors a201 and b202 are both hinged to the rotor arm 101 of the body system 100, and the cams a203 and b204 are respectively connected to the clamping plates a102 and b103 through bearings; the transmission system 300 includes a cam connecting rod assembly 301, a diagonal rod assembly 302, a diamond rod assembly 303, a curved rod 304, a support plate 305, and a cam connecting rod assembly concentric shaft 306. , the inclined rod assembly concentric axis 307, the middle concentric axis 308, and the crank rod concentric axis 309; one end of the cam connecting rod assembly 301 is hinged to the ends of the cam a203 and the cam b204 through the cam connecting rod assembly concentric axis 306, one end of the inclined rod assembly 302 is connected to the splint a102 and the splint b103 of the body system 100 through the inclined rod assembly concentric axis 307, and one end of the diamond rod assembly 303 is connected to the splint a102 and the splint b103 of the body system 100 through the middle concentric axis 308 is connected to the cam-link assembly 301 and the inclined rod assembly 302 by bearings, the diamond rod assembly 303 is hinged to the bent rod 304 at its middle by a bearing, the left end of the bent rod 304 is connected to the clamp a102 and the clamp b103 by the bent rod concentric axis 309, and the support plate 305 is hinged to the right end of the bent rod 304; the motor support 403 of the rotor system 400 is hinged to the diamond rod assembly 303 and the support plate 305 by bearings.

[0056] like Figure 5 As shown, the body system 100 includes a rotor arm 101, a splint a102, a splint b103, and a rotor arm skin 104; the splint a102 and the splint b103 are fixedly connected to the right end of the rotor arm 101, and are arranged symmetrically on the left and right, and the rotor arm skin 104 covers the outside of the rotor arm 101.

[0057] The rotor arm 101 is an I-shaped beam box structure. The right end face of the rotor arm 101 is provided with a tilt mechanism helicopter mode limit slope 1011 and a fixed-wing aircraft mode limit slope 1012. The left end of the rotor arm 101 is fixedly connected to the wing load-bearing structure.

[0058] The splint a102 includes a self-locking pin a1021 and an electromagnetic spring a1022; the self-locking pin a1021 and the electromagnetic spring a1022 are fixedly connected; the splint b103 contains a self-locking pin b1031 and an electromagnetic spring b1032, the self-locking pin b1031 and the electromagnetic spring b1032 are fixedly connected, and the self-locking pin a1021 and the self-locking pin b1031 are respectively arranged on the splint a102 and the splint b103 in a distributed circle.

[0059] like Figure 6 As shown, the linear motor a201 includes a motor base a2011 and a retractable push rod a2012; the retractable push rod a2012 is installed inside the motor base a2011, and the linear motor b202 includes a motor base b2021 and a retractable push rod b2022; the retractable push rod b2022 is installed inside the motor base b2021.

[0060] The cam a203 has a helicopter mode limiting slope a2031 and a fixed-wing aircraft mode limiting slope a2032, and the cam b204 has a helicopter mode limiting slope b2041 and a fixed-wing aircraft mode limiting slope b2042. In the helicopter mode, the helicopter mode limiting slope a2031 and the helicopter mode limiting slope b2041 contact the helicopter mode limiting slope 1011 of the rotor arm 101; in the fixed-wing aircraft mode, the fixed-wing aircraft mode limiting slope a2032 and the fixed-wing aircraft mode limiting slope b2042 contact the fixed-wing aircraft mode limiting slope 1012 of the rotor arm 101.

[0061] A circle of self-locking holes a2033 are distributed on the surface of cam a203, and a circle of self-locking holes b2043 are distributed on the surface of cam b204. In the event of an accidental power failure, in addition to the linear motor being self-locking due to power failure, the self-locking pin a1021 will be inserted into the self-locking hole a2033 under the action of the electromagnetic spring a1022; the self-locking pin b1031 will be inserted into the self-locking hole b2043 under the action of the electromagnetic spring b1032, and together with the linear motor, prevent the tilting mechanism from rotating at will.

[0062] Linear motor a201 and linear motor b202 have built-in dual Hall elements.

[0063] The sizes of cam a203 and cam b204 and the positions of their ears always meet the minimum force arm value requirements of linear motor a201 and linear motor b202 during the tilting process.

[0064] like Figure 7 As shown, the cam-link assembly 301 includes a cam link a3011 and a cam link b3012 which are symmetrically arranged on the left and right; one end of the cam link a3011 and the cam link b3012 are respectively hinged to the cam a203 and the cam b204 through the coaxial shaft 306 of the cam-link assembly, the other end of the cam link a3011 is hinged to the oblique rod a3021 and the diamond rod a3031 through the middle concentric shaft 308, and the other end of the cam link b3012 is hinged to the oblique rod b3022 and the diamond rod b3032 through the middle concentric shaft 308.

[0065] The diagonal tie rod assembly 302 includes a diagonal tie rod a3021 and a diagonal tie rod b3022 which are arranged symmetrically on the left and right. One end of the diagonal tie rod a3021 and the diagonal tie rod b3022 are connected to the coaxial axis 307 of the diagonal tie rod assembly by a bearing, and the other end is hinged to the cam connecting rod a3011, the diamond rod a3031, the cam connecting rod b3012 and the diamond rod b3032 through the middle concentric axis 308. The coaxial axis 307 of the diagonal tie rod assembly is connected to the splint a102 and the splint b103 for anti-rotation.

[0066] The diamond rod assembly 303 includes a diamond rod a3031 and a diamond rod b3032 which are symmetrically arranged on the left and right; the left end of the diamond rod a3031 is hinged to the cam connecting rod a3011 and the inclined rod a3021 through the middle concentric shaft 308, and the left end of the diamond rod b3032 is hinged to the cam connecting rod b3012 and the inclined rod b3022 through the middle concentric shaft 308. The diamond rod a3031 and the diamond rod b3032 are connected to both sides of the curved rod 304 through bearings in the middle.

[0067] The left end of the bent rod 304 is connected to the bent rod concentric shaft 309 through a bearing, and the right end is hinged to the support plate 305. The bent rod concentric shaft 309 and the clamping plates a102 and b103 are anti-rotationally connected.

[0068] like Figure 8 As shown, rotor system 400 includes fairing 401, rotor drive motor 402, motor support 403, rotor hub 404, and rotor blades 405; rotor drive motor 402 and motor support 403 are installed inside fairing 401, rotor drive motor 402 is connected to rotor hub 404, and rotor blades 405 are installed on rotor hub 404.

[0069] There are four groups of ears on the lower end surface of the motor support 403, including ear a4031, ear b4032, ear c4033 and ear d4034. Motor heat dissipation holes are distributed on the motor support 403. Ear a4031 and ear b4032 are hinged to the diamond rod b3032 and diamond rod a3031 respectively. Ear c4033 and ear d4034 are hinged to the support plate 305 through bearings.

[0070] The working principle of the rotor tilting mechanism of an electric vertical take-off and landing aircraft of the present invention is as follows:

[0071] like Figure 1(a) shows a schematic diagram of the helicopter mode of the rotor tilt mechanism. In this state, the linear motor a201 and the linear motor b202 are in the maximum extension state, the retractable push rod a2012 extends from the inside of the motor base a2011, and the retractable push rod b2022 extends from the inside of the motor base b2021. The internal electronic limit switches of the linear motor a201 and the linear motor b202 are locked at the extension limit position. The cam a203 and the cam b204 rotate counterclockwise to the maximum angle under the push of the retractable push rod a2012 and the retractable push rod b2022. Under the constraints of the geometric relationship and connection relationship of the various components of the transmission system 300, the cam connecting rod assembly 301, the inclined rod assembly 302, the diamond rod assembly 303, the curved rod 304 and the support plate 305 are in Figure 1 (a) shows the position, under the direct action of the diamond rod assembly 303 and the support plate 305, the motor support 403 is in a horizontal state, and the tilt mechanism can support the rotation of the rotor to perform vertical take-off and landing or hovering flight. Fig.11 As shown in (a), the helicopter mode limiting inclined surface a2031 on the cam a203 contacts the helicopter mode limiting inclined surface 1011 on the rotor arm 101, limiting the counterclockwise rotation of the cam a203; similarly, the helicopter mode limiting inclined surface b2041 on the cam b204 contacts the helicopter mode limiting inclined surface 1011 on the rotor arm 101, limiting the counterclockwise rotation of the cam b204, so that the rotor blade disc is accurately at the tilt angle of 90°. Fig. 9 As shown in (a), in helicopter mode, the rotor fairing 401 is located to the upper right of the rotor arm skin 104, and the two do not interfere with each other.

[0072] like Figure 1(b) shows a schematic diagram of the fixed-wing aircraft mode of the rotor tilt mechanism. When the tilt mechanism is converted from the helicopter mode to the fixed-wing aircraft mode, the linear motor a201 and the linear motor b202 gradually shrink from the extended state, and the retractable push rod a2012 and the retractable push rod b2022 shrink to drive the cam a203 and the cam b204 to rotate clockwise, thereby driving the transmission system 300 to move to the lower left as a whole. During the extension and retraction process, the linear motor a201 and the linear motor b202 will rotate around their installation points on the rotor arm 101, causing the arm length, force size and extension speed of the linear motor a201 and the linear motor b202 to change dynamically, ensuring that the tilt process is continuous and stable. During the tilt process, the built-in dual Hall feedback of the linear motor a201 and the linear motor b202 can provide real-time feedback on the tilt angle, which is convenient for the control system to make accurate judgments and operations. In the fixed-wing aircraft mode, the linear motor a201 and the linear motor b202 are in a retracted state, their internal electronic limit switches are locked in the retracted extreme position, the retractable push rod a2012 is retracted back into the motor base a2011, the retractable push rod b2022 is retracted back into the motor base b2021, and the cam-link assembly 301, the inclined rod assembly 302, the diamond rod assembly 303, the curved rod 304 and the support plate 305 of the transmission system 300 are in Figure 1 (b) shows the position, at this time, the above-mentioned rods are contracted in a small area, and there is no interference between the rods. Under the pulling action of the diamond rod assembly 303 and the support plate 305, the motor support 403 is in a vertical state. At this time, the tilt mechanism can support the electric vertical take-off and landing aircraft to fly forward at high speed. Fig.11 As shown in (b), the fixed-wing aircraft mode limiting inclined surface a2032 on the cam a203 contacts the fixed-wing aircraft mode limiting inclined surface 1012 on the rotor arm 101, limiting the clockwise rotation of the cam a203; similarly, the fixed-wing aircraft mode limiting inclined surface b2042 on the cam b204 contacts the fixed-wing aircraft mode limiting inclined surface 1012 on the rotor arm 101, limiting the clockwise rotation of the cam b204, so that the rotor blade disc is accurately at the tilt angle 0° position. Fig. 9 As shown in (b), the rotor fairing 401 is in contact and closed with the rotor arm skin 104 to form a complete aerodynamic shape, thereby reducing the adverse effect of the tilt mechanism on the high-speed forward flight performance.

[0073] When switching from fixed-wing aircraft mode to helicopter mode, the retractable push rod a2012 gradually extends from the inside of the motor base a2011, and the retractable push rod b2022 gradually extends from the inside of the motor base b2021, pushing the cam a203 and the cam b204 to rotate counterclockwise, thereby driving the transmission system 300 to move to the upper right. Under the action of the transmission system 300, the motor support 403 tilts from a vertical state to a horizontal state until it reaches the 90° tilt position corresponding to the helicopter mode. At this time, the helicopter mode limit bevel a2031 on the cam a203 and the helicopter mode limit bevel b2041 on the cam b204 contact the helicopter mode limit bevel 1011 on the rotor arm 101, limiting excessive tilting of the cam a203 and the cam b204.

[0074] like Fig.12 As shown in (a), the mechanical self-locking system between the clamp a102 and the cam a203 is used as an example for explanation. When the linear motor a201 and the linear motor b202 are both powered on, the electromagnetic spring a1022 distributed on the clamp a102 is compressed by the electromagnetic force, which constrains the self-locking pin a1021 to be within the plane of the clamp a102 and will not affect the normal rotation of the cam a203. If the linear motor a201 and the linear motor b202 are both powered off, on the one hand, the linear motor a201 and the linear motor b202 will be powered off and self-locked, and on the other hand, the mechanical self-locking system of the tilting mechanism will be activated, such as Fig.12 As shown in (b), the electromagnetic spring a1022 is not affected by the electromagnetic force after power failure, and changes from a compressed state to an extended state. The self-locking pin a1021 is inserted into the self-locking hole a2033 under the elastic action of the electromagnetic spring a1022, limiting the rotation of the cam a203, so that the flight control system can take timely measures to ensure flight safety. The mechanical self-locking system between the clamp b103 and the cam b204 is the same as the above-mentioned mechanical self-locking system. Fig.10 As shown, by designing the installation points of the linear motor a201 and the linear motor b202 on the rotor arm 101 to be concentric, the rotor tilt mechanism arranged symmetrically on both sides is connected as a whole by combining the cam connecting rod assembly concentric shaft 306, the inclined rod assembly concentric shaft 307, the middle concentric shaft 308, the crankshaft concentric shaft 309 and other components, which can effectively improve the stability and vibration resistance of the tilt mechanism. In addition, under the above-mentioned mechanism arrangement, when one of the linear motors a201 / linear motors b202 fails, the entire tilt mechanism can be driven to continue working by increasing the instantaneous power of the other linear motor b202 / linear motor a201, thereby improving flight safety through power redundancy design.

Claims

1. An electric vertical take-off and landing aircraft rotor tilting mechanism, characterized in that: The invention comprises a body system (100), a drive system (200), a transmission system (300), and a rotor system (400); the drive system (200) comprises a linear motor a (201), a linear motor b (202), a cam a (203), and a cam b (204); the front ends of the linear motor a (201) and the linear motor b (202) are respectively connected to the ear pieces of the cam a (203) and the cam b (204) through bearings; the linear motor a (201) and the linear motor b The end mounting points of (202) are hinged to the rotor arm (101) of the body system (100); the cam a (203) and the cam b (204) are respectively connected to the clamping plate a (102) and the clamping plate b (103) through bearings; the transmission system (300) includes a cam connecting rod assembly (301), a diagonal tie rod assembly (302), a diamond rod assembly (303), a curved rod (304), a support plate (305), a cam connecting rod assembly concentric shaft (306), and a diagonal tie rod assembly concentric shaft (307), a middle concentric shaft (308), and a crankshaft concentric shaft (309); one end of the cam connecting rod assembly (301) is hinged to the ends of cam a (203) and cam b (204) through the cam connecting rod assembly concentric shaft (306); one end of the inclined rod assembly (302) is connected to the clamping plate a (102) and clamping plate b (103) of the body system (100) through the inclined rod assembly concentric shaft (307); one end of the rhombus rod assembly (303) is connected to the clamping plate a (102) and clamping plate b (103) of the body system (100) through the middle concentric shaft (308). The cam connecting rod assembly (301) and the inclined rod assembly (302) are connected by bearings, the diamond rod assembly (303) is hinged to the bent rod (304) at its middle through a bearing, the left end of the bent rod 304 is connected to the clamping plate a (102) and the clamping plate b (103) through the bent rod concentric axis (309), and the support plate (305) is hinged to the right end of the bent rod (304); the motor support (403) of the rotor system (400) is hinged to the diamond rod assembly (303) and the support plate (305) through bearings.

2. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The body system (100) comprises a rotor arm (101), a clamping plate a (102), a clamping plate b (103), and a rotor arm skin (104); the clamping plate a (102) and the clamping plate b (103) are fixedly connected to the right end of the rotor arm (101) and arranged symmetrically on the left and right sides; the rotor arm skin (104) covers the outer side of the rotor arm (101).

3. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The rotor arm (101) is an I-shaped beam box structure, the right end surface of the rotor arm (101) is provided with a tilt mechanism helicopter mode limiting inclined surface (1011) and a fixed-wing aircraft mode limiting inclined surface (1012), and the left end of the rotor arm (101) is fixedly connected to the wing load-bearing structure.

4. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The clamping plate a (102) includes a self-locking pin a (1021) and an electromagnetic spring a (1022); the self-locking pin a (1021) and the electromagnetic spring a (1022) are fixedly connected; the clamping plate b (103) includes a self-locking pin b (1031) and an electromagnetic spring b (1032); the self-locking pin b (1031) and the electromagnetic spring b (1032) are fixedly connected, and the self-locking pin a (1021) and the self-locking pin b (1031) are respectively arranged on the clamping plate a (102) and the clamping plate b (103) in a distributed circle.

5. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The linear motor a (201) and the linear motor b (202) have dual Hall elements built in.

6. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The sizes of the cam a (203) and the cam b (204) and the positions of their lugs always meet the minimum force arm value requirements of the linear motor a (201) and the linear motor b (202) during the tilting process.

7. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The linear motor a (201) comprises a motor base a (2011) and a retractable push rod a (2012); the retractable push rod a (2012) is installed inside the motor base a (2011); the linear motor b (202) comprises a motor base b (2021) and a retractable push rod b (2022); the retractable push rod b (2022) is installed inside the motor base b (2021).

8. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The cam a (203) has a helicopter mode limiting inclined surface a (2031) and a fixed-wing aircraft mode limiting inclined surface a (2032); the cam b (204) has a helicopter mode limiting inclined surface b (2041) and a fixed-wing aircraft mode limiting inclined surface b (2042); in the helicopter mode, the helicopter mode limiting inclined surface a (2031) and the helicopter mode limiting inclined surface b (2041) are in contact with the helicopter mode limiting inclined surface (1011) of the rotor arm (101); in the fixed-wing aircraft mode, the fixed-wing aircraft mode limiting inclined surface a (2032) and the fixed-wing aircraft mode limiting inclined surface b (2042) are in contact with the fixed-wing aircraft mode limiting inclined surface (1012) of the rotor arm (101).

9. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: A circle of self-locking holes a (2033) are distributed on the surface of the cam a (203), and a circle of self-locking holes b (2043) are distributed on the surface of the cam b (204). When the power is accidentally cut off, in addition to the linear motor being self-locked when the power is cut off, the self-locking pin a (1021) will be inserted into the self-locking hole a (2033) under the action of the electromagnetic spring a (1022); the self-locking pin b (1031) will be inserted into the self-locking hole b (2043) under the action of the electromagnetic spring b (1032), and together with the linear motor, prevent the tilting mechanism from rotating at will.

10. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The cam link assembly (301) comprises a cam link a (3011) and a cam link b (3012) which are arranged symmetrically on the left and right. One end of the cam link a (3011) and the cam link b (3012) are respectively hinged to the cam a (203) and the cam b (204) through a cam link assembly coaxial shaft (306), the other end of the cam link a (3011) is hinged to the inclined rod a (3021) and the diamond rod a (3031) through a middle coaxial shaft (308), and the other end of the cam link b (3012) is hinged to the inclined rod b (3022) and the diamond rod b (3032) through a middle coaxial shaft (308).

11. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The inclined rod assembly (302) comprises an inclined rod a (3021) and an inclined rod b (3022) which are arranged symmetrically on the left and right. One end of the inclined rod a (3021) and the inclined rod b (3022) are connected to the coaxial axis (307) of the inclined rod assembly by a bearing, and the other end is hinged to the cam connecting rod a (3011), the rhombus rod a (3031), the cam connecting rod b (3012) and the rhombus rod b (3032) through the middle coaxial axis (308). The coaxial axis (307) of the inclined rod assembly is connected to the clamping plate a (102) and the clamping plate b (103) for anti-rotation.

12. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The diamond rod assembly (303) comprises a diamond rod a (3031) and a diamond rod b (3032) which are arranged symmetrically on the left and right. The left end of the diamond rod a (3031) is hinged to the cam connecting rod a (3011) and the inclined rod a (3021) through a central concentric shaft (308), and the left end of the diamond rod b (3032) is hinged to the cam connecting rod b (3012) and the inclined rod b (3022) through a central concentric shaft (308). The diamond rod a (3031) and the diamond rod b (3032) are connected to both sides of the curved rod (304) through bearings in the middle.

13. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The left end of the curved rod (304) is connected to the curved rod coaxial shaft (309) through a bearing, and the right end is hinged to the support plate (305). The curved rod coaxial shaft (309) and the clamping plates a (102) and b (103) are connected to prevent rotation.

14. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The rotor system (400) comprises a fairing (401), a rotor drive motor (402), a motor support (403), a rotor hub (404), and rotor blades (405); the rotor drive motor (402) and the motor support (403) are installed inside the fairing (401), the rotor drive motor (402) is connected to the rotor hub (404), and the rotor blades (405) are installed on the rotor hub (404).

15. The electric vertical take-off and landing aircraft rotor tilting mechanism according to claim 1, characterized in that: The lower end surface of the motor support (403) is provided with four groups of lugs, including lug a (4031), lug b (4032), lug c (4033) and lug d (4034). Motor heat dissipation holes are distributed on the motor support (403). Lug a (4031) and lug b (4032) are hinged to rhombus rod b (3032) and rhombus rod a (3031) respectively. Lug c (4033) and lug d (4034) are hinged to the support plate (305) through bearings.

16. A method for operating a rotor tilt mechanism of an electric vertical take-off and landing aircraft according to any one of claims 1 to 15, characterized in that: After receiving instructions from the flight control system, the linear motor a (201) and the linear motor b (202) drive the retractable push rod a (2012) and the retractable push rod b (2022) to retract, thereby driving the cam a (203) and the cam b (204) to rotate. The rotation of the cam a (203) and the cam b (204) causes the transmission system (300) to move and rotate synchronously. The transmission system (300) gradually transmits the input of the drive system (200) to the output end component of the transmission system (300), and finally realizes the conversion of the rotor system (400) between the helicopter mode and the fixed-wing aircraft mode. When the power is on, the electromagnetic spring a (1022) and the electromagnetic spring b (1032) are in a compressed state due to the magnetic force, pulling the self-locking pin a (1021) and the self-locking pin b (1031) to shrink within the plane of the clamping plate a (102) and the clamping plate b (103); when the power is off, the electromagnetic spring a (1022) and the electromagnetic spring b (1032) lose the magnetic force restriction and are in an extended state, and the self-locking pin a (1021) and the self-locking pin b (1031) exceed the plane of the clamping plate a (102) and the clamping plate b (103); If the linear motor a (201) or the linear motor b (202) is accidentally powered off during the tilting process, the instantaneous power of the linear motor b (202) or the linear motor a (201) is increased to keep the tilting motion going; if the linear motor a (201) and the linear motor b (202) are accidentally powered off at the same time, the self-locking pin a (1021) will be inserted into the self-locking hole a (2033) under the action of the electromagnetic spring a (1022); the self-locking pin b (1031) will be inserted into the self-locking hole b (2043) under the action of the electromagnetic spring b (1032), further enhancing the self-locking effect on the basis of the self-locking function of the linear motor; In helicopter mode, the helicopter mode limit slope a (2031) and the limit slope b (2041) are in contact with the helicopter mode limit slope (1011) of the rotor arm (101); in fixed-wing aircraft mode, the fixed-wing aircraft mode limit slope a (2032) and the limit slope b (2042) are in contact with the fixed-wing aircraft mode limit slope (1012) of the rotor arm (101). In addition to the position control of the linear motor a (201) and the linear motor b (202), the above-mentioned method further limits excessive tilting of the tilt mechanism at two extreme positions.

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