CN113753231A

CN113753231A - Aircraft and coaxial dual-rotor assembly

Info

Publication number: CN113753231A
Application number: CN202111183897.2A
Authority: CN
Inventors: 王谭; 梁毅诚; 娄津源; 史翊辰; 曹雪宇; 王伟民
Original assignee: Guangdong Huitian Aerospace Technology Co Ltd
Current assignee: Guangdong Huitian Aerospace Technology Co Ltd
Priority date: 2021-10-11
Filing date: 2021-10-11
Publication date: 2021-12-07
Also published as: WO2023060678A1

Abstract

The invention provides an aircraft, which comprises a fuselage, a wing assembly, a tilt rotor assembly and a coaxial dual-rotor assembly, wherein the wing assembly is connected to the fuselage; the tilt rotor assembly is tiltably connected to the wing assembly; coaxial dual rotor subassembly includes installation axle, first paddle and second paddle, and the erection joint is in the fuselage, and first paddle and second paddle rotationally connect respectively in the relative both ends of installation axle, and the rotation direction of first paddle is opposite with the rotation direction of second paddle. According to the aircraft provided by the invention, the coaxial double-rotor assembly is arranged, and the rotating direction of the first blade of the coaxial double-rotor assembly is opposite to that of the second blade, so that the reaction torque generated by the aircraft can be counteracted, the course stability of the aircraft is improved, and the flight safety of the aircraft is ensured. The invention also provides a coaxial dual-rotor assembly.

Description

Aircraft and coaxial dual-rotor assembly

Technical Field

The invention relates to the technical field of flight equipment, in particular to an aircraft and a coaxial double-rotor-wing assembly.

Background

Along with the continuous development of science and technology, the application of aircraft is also more and more extensive, and the aircraft can also be used to people's daily trip except being used for video shooting, agricultural watering and fire control rescue.

However, in the flight process of the existing aircraft, reaction torque is generated by rotation of the propeller, so that the aircraft has a course deviation phenomenon, and a great flight hidden trouble is brought.

Disclosure of Invention

It is an object of embodiments of the present invention to provide an aircraft and a coaxial dual rotor assembly that improve upon the above-mentioned problems. The embodiment of the invention achieves the aim through the following technical scheme.

In a first aspect, the present invention provides an aircraft comprising a fuselage, a wing assembly, a tiltrotor assembly, and a coaxial dual-rotor assembly, the wing assembly being connected to the fuselage; the tilt rotor assembly is tiltably connected to the wing assembly; coaxial dual rotor subassembly includes installation axle, first paddle and second paddle, and installation hub connection is in the fuselage, and first paddle and second paddle rotationally connect respectively in the relative both ends of installation axle, and the direction of rotation of first paddle is opposite with the direction of rotation of second paddle, can offset the counter torque that the aircraft produced, has promoted the stability in aircraft's course, has guaranteed the flight safety of aircraft.

In an embodiment, the fuselage is provided with an accommodating cavity and an opening, the opening is communicated with the accommodating cavity, and the coaxial dual-rotor assembly is rotatably arranged in the accommodating cavity and exposed out of the opening, so that the resistance of the aircraft in the flying process is reduced, and the high-speed flat flying of the aircraft is facilitated.

In one embodiment, the fuselage further comprises a cover for covering the opening, further reducing drag experienced by the aircraft during flight.

In an embodiment, the fuselage further comprises a main body part, the main body part is used for arranging the wing assembly, and the cover plate is rotatably arranged on the main body part, so that the cover plate is simple in connection with the main body part, and good in connection stability.

In one embodiment, the wing assembly includes a fixed portion disposed on the fuselage and a flap portion movably connected to the fixed portion and foldable relative to the fixed portion to enable the wing assembly to be folded to reduce the overall size of the aircraft, so that the aircraft can be parked in a road transition or a narrow parking space while still having the possibility of parking and transferring, thereby improving the flexibility of the aircraft during parking and transferring.

In one embodiment, the fixed portion extends along a first direction, the folding portion is provided with a first rotating axis and a second rotating axis, the first rotating axis extends along the first direction, the second rotating axis extends along a second direction, the first blade and the second blade are sequentially arranged along a third direction, and the first direction, the second direction and the third direction are perpendicular to each other in pairs, so that the wing assembly can be folded step by step, and damage to the wing assembly in the folding process is reduced.

In one embodiment, the aircraft further includes a folding mechanism disposed on the fixing portion, and the folding portion can be folded relative to the fixing portion through the folding mechanism, so that the folding relative to the fixing portion is smoother, and the folding efficiency of the folding portion is improved.

In one embodiment, the folding mechanism comprises a first rotating shaft and a second rotating shaft, the folding part rotates along the first rotating axis through the first rotating shaft, and the folding part rotates along the second rotating axis through the second rotating shaft, so that the wing assembly can be folded step by step, and damage to the wing assembly in the folding process is reduced.

In an embodiment, the section of turning over includes first section of turning over and the second section of turning over, and first section of turning over movably connects in the fixed part to relative fixed part can be turned over, and the rotor assembly that verts sets up in the second section of turning over, and the rotor assembly that verts sets up in the one end that the fuselage was kept away from to the wing assembly, avoids the wing assembly to the rotor assembly that verts to produce the influence.

In an embodiment, the tilting rotor assembly and the second turning section form a power mechanism, and the gravity center of the power mechanism is located on the rotation axis of the second turning section, so that the gravity center of the power mechanism cannot be changed by the rotation of the tilting rotor assembly, the stability of the gravity center of the power mechanism can be ensured, and the efficiency of the vertical lifting force is ensured.

In one embodiment, the length of the fixing part is 10% -30% of the length of the wing assembly, so that the width of the folded aircraft is greatly reduced, and transition of the aircraft is facilitated.

In an embodiment, the aircraft still includes the connecting piece that verts, and the connecting piece that verts rotationally sets up in the wing subassembly, and the rotor subassembly that verts is connected with the connecting piece transmission that verts to change space angle under the drive of the connecting piece that verts, with the switching between the rotor state that realizes the rotor subassembly that verts and the stationary vane state.

In an embodiment, the rotor subassembly that verts includes the rotor body and the rotor mount pad that verts, and the rotor mount pad is connected with the connecting piece transmission that verts, and the rotor body that verts rotationally sets up in the rotor mount pad, and the rotor body that verts includes the rotor paddle that verts, and the rotor paddle that verts can buckle relative to the rotor mount pad, can reduce the length of aircraft, has improved the flexibility of aircraft at berth and transition in-process.

In an embodiment, the wing subassembly is two, and two wing subassemblies are connected respectively in the relative both sides of fuselage, and every wing subassembly all is equipped with a rotor subassembly that verts, and two rotor subassemblies that vert set up about the focus symmetry of aircraft for the rotor subassembly that verts can not change the position of the focus of aircraft at the in-process that verts, has increased the stability of aircraft at the flight in-process.

In an embodiment, four rotor modules are constituteed with coaxial two rotor assemblies to two tilt rotor assemblies, and the focus of aircraft is located four rotor modules's center for the focus's of aircraft position does not receive the influence of coaxial two rotor assemblies's state and the state that tilt rotor assembly located, strengthens the aircraft stability at the flight in-process.

In one embodiment, the center of the coaxial dual rotor assembly is spaced from the center of gravity of the aircraft by a distance greater than the distance between the center of gravity of the tiltrotor assembly and the aircraft such that the overall primary lift source of the aircraft is centered on the tiltrotor assembly.

In one embodiment, the aircraft further comprises a tail wing, wherein the tail wing is connected to the fuselage to avoid wake interference and improve the control efficiency of the horizontal tail.

In one embodiment, the tail wing includes a horizontal wing and a vertical wing connected between the fuselage and the horizontal wing and extending upwardly, the horizontal wing extending in the same direction as the wing assembly to control the attitude of the aircraft.

In one embodiment, the aircraft further comprises wheel sets, the wheel sets are arranged at the bottom of the fuselage, and the fuselage is movable along the ground through the wheel sets, so that the transition requirement of the aircraft is met.

In a second aspect, the present invention further provides a coaxial dual-rotor assembly, which includes a mounting shaft, a first blade and a second blade, wherein the first blade and the second blade are rotatably connected to two opposite ends of the mounting shaft, and a rotation direction of the first blade is opposite to a rotation direction of the second blade, so that a counter torque generated by rotation of the first blade and a counter torque generated by rotation of the second blade can be cancelled out.

Compared with the prior art, the aircraft and the coaxial dual-rotor assembly provided by the invention have the advantages that the rotating direction of the first blade of the coaxial dual-rotor assembly is opposite to that of the second blade, so that the counter torque generated by the aircraft can be offset, the course stability of the aircraft is improved, and the flight safety of the aircraft is ensured.

These and other aspects of the invention are apparent from and will be elucidated with reference to the embodiments described hereinafter.

Drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present application, the drawings needed to be used in the description of the embodiments are briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present application, and it is obvious for those skilled in the art to obtain other drawings based on these drawings without creative efforts.

Fig. 1 is a schematic structural diagram of an aircraft provided by an embodiment of the invention.

FIG. 2 is a schematic illustration of the aircraft of FIG. 1 in a VTOL condition.

Fig. 3 is a schematic structural view of the aircraft shown in fig. 2 from another perspective.

Fig. 4 is a schematic structural view of a tilt connection and rotor assembly of the aircraft shown in fig. 1.

FIG. 5 is a schematic illustration of the aircraft shown in FIG. 1 in a level flight cruise condition.

FIG. 6 is a schematic illustration of the configuration of the aircraft shown in FIG. 1 during a ground transition condition.

Fig. 7 is a schematic structural view of the aircraft shown in fig. 6 from another perspective.

Fig. 8 is a schematic structural view of a coaxial dual rotor assembly provided by an embodiment of the present invention.

Detailed Description

In order to facilitate an understanding of the embodiments of the present invention, the embodiments of the present invention will be described more fully hereinafter with reference to the accompanying drawings. Preferred embodiments of the present invention are shown in the drawings. This invention may, however, be embodied in many different forms and should not be construed as limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete.

Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. The terminology used herein in the examples of the present invention is for the purpose of describing particular embodiments only and is not intended to be limiting of the invention.

The inventor of the application discovers through research that various urban aircrafts are developed by a plurality of companies at present, including urban helicopters, aircrafts adopting a multi-rotor principle, aircrafts adopting short-distance running and landing and aircrafts adopting multi-tilting rotors.

However, existing aircraft designs do not compromise on-road travelMost of the functions of driving, vertical taking off and landing, horizontal flying and transition cannot drive on the road or take off vertically. Such as Yihang 216 and

the self is only provided with a fixed undercarriage, the self cannot move autonomously after landing, and passengers need to get off the airplane and then arrive at a destination by adopting other traffic modes. Another example is Joby

S4, a six-rotor tilt rotor scheme is adopted, the wingspan is large, and the six-rotor tilt rotor can only be lifted and landed on a fixed special parking apron and cannot meet the requirements of transition. For example

The mode of combining a tricycle and a gyroplane is adopted, so that the running is needed for taking off and landing, and the taking off and landing can be carried out only by a barrier-free runway with a certain distance. Also for example Klein

The mode of combining the fixed wings and the automobile is adopted, the takeoff and the landing need to be run, and the takeoff and the landing need to be carried out only on a barrier-free runway with a certain distance.

The tilting rotor aircraft with the traditional double rotors has the characteristics of a helicopter and a fixed wing. Compared with a fixed-wing aircraft, the tilt rotor wing can take off and land vertically without depending on an airport runway; compared with the traditional helicopter, the tilting rotor wing has larger cruising speed and range, and flies in the form of a fixed wing during cruising, so that the tilting rotor wing is more economical than the helicopter.

The existing tilting rotor type comprises V22 and V280 of Bell company, which adopt the conventional layout of a transverse double rotor, depend on the same rotor cycle variable pitch control attitude as a helicopter in the conversion stage of vertical flight and vertical plane flight, and mainly adopt the same control plane control attitude as a fixed wing in the plane flight configuration.

The existing dual-rotor tilt rotor type depends on a periodic variable pitch mechanism to realize control, when the longitudinal attitude of the aircraft is controlled by using the periodic variable pitch mechanism, the coupling effect of forward flight of the aircraft can be caused, the periodic variable pitch mechanism has a complex structure, the complex structure and the control mode greatly increase the research and development and manufacturing cost of the aircraft, and meanwhile, great challenges are brought to safety and reliability.

The existing aircraft product does not perfectly solve the problems of long range, vertical landing and road surface transition, and the aircraft for the future cities and the intercity has the functions to bring a brand-new aircraft with high safety to the market so as to meet the market demand.

The invention aims to provide an aircraft and a coaxial double-rotor assembly aiming at the defects of the existing aircraft technology, which can counteract the reaction torque generated by the aircraft, improve the course stability of the aircraft, ensure the flight safety of the aircraft, have the functions of vertical take-off and landing, flat flight and transition and have wide application prospect.

The invention provides an aircraft and a coaxial dual-rotor assembly, which are described in detail in the following with reference to the detailed description and the attached drawings.

Referring to fig. 1 and 2, the present invention provides an aircraft 1 comprising a fuselage 10, a wing assembly 20, a tiltrotor assembly 50, and a coaxial dual-rotor assembly 60. The wing assembly 20 is connected to the fuselage 10; tiltrotor assembly 50 is tiltably connected to wing assembly 20; the coaxial dual-rotor assembly 60 includes a mounting shaft 61, a first blade 63 and a second blade 65, the mounting shaft 61 is connected to the fuselage 10, the first blade 63 and the second blade 65 are rotatably connected to opposite ends of the mounting shaft 61, respectively, and the first blade 63 rotates in a direction opposite to that of the second blade 65.

In the present embodiment, the aircraft 1 may be an aircraft for cities, i.e. an aircraft that performs traffic replenishment between cities. In other embodiments, the aircraft 1 may also be used in the fields of video filming, agricultural irrigation, fire rescue, and the like.

The body 10 is provided with a receiving cavity 11 and an opening 13, and the opening 13 is communicated with the receiving cavity 11. In the present embodiment, the receiving cavity 11 is provided substantially at the rear portion of the fuselage 10 for receiving the coaxial dual-rotor assembly 60. In other embodiments, the receiving cavity 11 may be disposed at other positions of the body 10, for example, at the middle of the body 10. The opening 13 may be used to expose the coaxial dual-rotor assembly 60 and allow the first and

second blades

63, 65 of the coaxial dual-rotor assembly 60 to extend out of the fuselage 10 during rotation. In the present embodiment, the number of the openings 13 is two, and the two openings 13 are corresponding in position and are respectively disposed on two opposite sides of the body 10. In other embodiments, the number of openings 13 may also be one.

Referring to fig. 2 and 3, in one embodiment, the fuselage 10 further includes a main body portion 15 and a cover 17, the main body portion 15 being used to provide the wing assembly 20, and the cover 17 being used to cover the opening 13 to further reduce drag experienced by the aircraft 1 during flight. In some embodiments, the cover plate 17 is rotatably disposed on the main body 15, so that the cover plate 17 can be connected to the main body 15 in a simple manner and with good connection stability, for example, by being hinged to the main body 15. In other embodiments, the covering plate 17 may be slidably engaged with the main body 15 to cover the opening 13 in a sliding manner, for example, the main body 15 may be provided with a guide rail, and the covering plate 17 is provided with a sliding slot engaged with the guide rail. In other embodiments, the cover plate 17 and the main body 15 may have other matching manners, so as to satisfy the purpose that the cover plate 17 can cover the opening 13.

The cross section of the wing assembly 20 is approximately oval, the wing assembly 20 can be used for bearing aerodynamic force, and the approximately oval structure of the wing assembly 20 enables air above the wing assembly 20 to have high flow speed and low pressure, and air below the wing assembly 20 to have low flow speed and high pressure, so that pressure difference is formed between the upper surface and the lower surface of the wing assembly 20, lifting force is generated on the fuselage 10, and the aircraft 1 can rise conveniently. The wing component 20 is in direct contact with the outside, and therefore the material of the wing component 20 needs to have high strength, good plasticity, smooth surface, and high corrosion resistance. The number of the wing assemblies 20 is two, the two wing assemblies 20 are respectively connected to two opposite sides of the fuselage 10, and the two wing assemblies 20 can also be connected with each other. The wing assembly 20 extends in a direction perpendicular to the extension of the fuselage 10.

Referring to fig. 1, the wing assembly 20 includes a fixing portion 21 and a folding portion 23, the fixing portion 21 and the folding portion 23 are sequentially disposed along an extending direction of the wing assembly 20, wherein the fixing portion 21 is disposed on the fuselage 10, for example, the fixing portion 21 may be fixed to the fuselage 10 by welding, riveting, or the like. The folding portion 23 is movably connected to the fixing portion 21 and can be folded relative to the fixing portion 21. The folding part 23 can be folded relative to the fixing part 21, so that the wing assembly 20 can be folded, the folding of the wing assembly 20 can reduce the overall size of the aircraft 1, particularly the width of the aircraft 1, so that the aircraft 1 can still have the possibility of parking and transferring when being parked at a road transition or a narrower parking position, and the flexibility of the aircraft 1 in the parking and transferring process is improved.

In this embodiment, the fixed portion 21 extends in a first direction, as do the wing assembly 20 and the flap portion 23. The length of the fixing portion 21 is approximately 10% -30% of the length of the wing assembly 20, for example, the length of the fixing portion 21 is approximately 15% of the length of the wing assembly 20, that is, the length of the turning portion 23 is approximately 85% of the length of the wing assembly 20, which greatly reduces the width of the folded aircraft 1 and facilitates the turning of the aircraft 1. In other embodiments, the fixed portion 21 may also be of other lengths, for example, the length of the fixed portion 21 is approximately 25% of the length of the wing assembly 20.

In this embodiment, the turning part 23 has a first axis of rotation extending in a first direction and a second axis of rotation extending in a second direction, the first direction being perpendicular to the second direction, so as to achieve stepwise folding of the wing assembly 20 and reduce damage to the wing assembly 20 during the folding process. In the present embodiment, the folding portion 23 is first rotated 90 ° upward or downward along the second rotation axis, and then rotated 90 ° backward along the first rotation axis. In other embodiments, the turning part 23 has a first rotation axis and a third rotation axis, wherein the third rotation axis extends along a third direction, the first direction, the second direction and the third direction are perpendicular to each other, and the turning part 23 can rotate 90 ° along the first rotation axis first and then rotate 90 ° backward along the third rotation axis. Where "up", "down" and "aft" are defined in terms of the perspective of normal flight of the aircraft 1.

In this embodiment, the wing assembly 20 extends in a first direction and the fuselage 10 extends in a second direction. For convenience of description, the first direction is defined as an X direction, the second direction is defined as a Y direction, the third direction is defined as a Z direction, and the three directions are perpendicular to each other.

The folding portion 23 includes a first folding section 231 and a second folding section 232, and the second folding section 232 is rotatable relative to the first folding section 231. When the folded portion 23 is not folded with respect to the fixed portion 21, that is, the folded portion 23 extends in the X direction, the rotation axis of the second folded section 232 extends in the X direction.

The first folding section 231 is movably connected to the fixing portion 21 and can be folded relative to the fixing portion 21, and the first folding section 231 can drive the second folding section 232 to fold. When the transition requirement is not met, the first folding section 231 is fixedly connected with the fixing portion 21, and when the first folding section 231 needs to rotate, the first folding section 231 can be manually adjusted or electrically controlled to be folded relative to the fixing portion 21, so that the wing assembly 20 can be folded, and the ground transition of the aircraft 1 is facilitated.

Referring to fig. 1 and 4, the aircraft 1 further includes a tilt connector 30, and the tilt connector 30 can be used to tilt the tilt rotor assembly 50 to change the state of the tilt rotor assembly 50. For example, the tilting connection member 30 is rotatably disposed at the first folding section 231, and the second folding section 232 is connected to the tilting connection member 30, so that the tilting connection member 30 can drive the second folding section 232 to rotate, so that the second folding section 232 rotates relative to the first folding section 231. Section 232 is rolled over to the second is equipped with tilt rotor assembly 50, and tilt rotor assembly 50 can be fixed to be set up in section 232 is rolled over to the second, for example, tilt rotor assembly 50 can be through bolted connection's mode and section 232 fixed connection is rolled over to the second. Can drive the second when setting up in the relative first section 231 that turns over of the connecting piece that turns over of the connecting piece 30 that turns over in first section 231 rotates and turn over section 232 rotation, the second turns over the rotation of section 232 and can drive the rotor subassembly 50 that turns over and vert.

The tilt linkage 30 is rotatably disposed on the wing assembly 20 (fig. 1), and the tilt linkage 30 is driven by a motor (not shown) to rotate.

In the present embodiment, the number of the tilt connectors 30 is two, and the two tilt connectors 30 are respectively disposed at the two first folding sections 231. For example, first section 231 of turning over is equipped with the motor, and the motor is connected with the transmission of the connecting piece 30 that verts, and the rotation of motor can drive the connecting piece 30 that verts and rotate to drive the rotor subassembly 50 that verts, change the state of rotor subassembly 50 that verts. In the present embodiment, the tilt connection 30 may be a hinge structure.

Tiltrotor assembly 50 is tiltably coupled to wing assembly 20 and has a rotor-wing state in which the axis of tiltrotor assembly 50 extends in the Z-direction and is capable of generating lift, and a fixed-wing state in which aircraft 1 is capable of vertical takeoff and landing; in the fixed-wing state, the axis of tiltrotor assembly 50 extends in the Y direction, and thrust can be generated, so that aircraft 1 has the capability of high-speed flat flight. By controlling the state of tiltrotor assembly 50, the flight mode of aircraft 1 can be controlled, and wing assembly 20 can be folded, enabling aircraft 1 to be parked and transferred flexibly. The aircraft 1 has the advantages of good flexibility and the like in the vertical take-off and landing, flat flight, mooring and transition processes, can meet the requirements of the aerial manned flight tasks in short and medium distances, and improves the practicability of the aircraft 1.

Tilt rotor subassembly 50 is connected with the transmission of connecting piece 30 that verts to change space angle under the drive of connecting piece 30 that verts, thereby realize tilting rotor subassembly 50's rotor state and the switching between the stationary vane state. For example, the connecting piece 30 that verts is connected through turning section 232 with the second, and the section 232 that turns over that drives the second verts to the drive sets up and turns over in the rotor assembly 50 that verts on the section 232 that turns over the second, guarantees that aircraft 1 is at VTOL and VTOL to the flat condition of flying the conversion under, and the thrust direction of rotor assembly 50 that verts can become upwards or forward, in order to satisfy the purpose that provides lift or thrust.

Referring to fig. 4 and 5, the tilt rotor assembly 50 includes a tilt rotor body 51 and a rotor mount 53, the rotor mount 53 is in transmission connection with the tilt connector 30, the tilt rotor body 51 is rotatably disposed on the rotor mount 53, the tilt rotor body 51 includes a tilt rotor blade 512, and the tilt rotor blade 512 is bendable relative to the rotor mount 53, so as to reduce the length of the aircraft 1 and improve the flexibility of the aircraft 1 during parking and transition. In addition, tiltrotor assembly 50 further includes a connection post 54, and rotor mount 53 is drivingly connected to tiltrotor connection 30 via connection post 54.

Tilt rotor assembly 50 sets up in the second section 232 of turning over, i.e., tilt rotor assembly 50 sets up in the one end that fuselage 10 was kept away from to wing assembly 20, avoids wing assembly 20 to tilt rotor assembly 50 and produces the influence. For example, tiltrotor assembly 50 may be fixedly connected to an end of second turning section 232 facing the direction of travel of aircraft 1 by attachment post 54. When the folded portion 23 extends along the X direction and the second folded section 232 does not rotate relative to the first folded section 231, the rotation axis of the tiltrotor assembly 50 extends along the Y direction, and the tiltrotor assembly 50 is at a horizontal position; when the folded portion 23 extends in the X direction and the second folded section 232 rotates relative to the first folded section 231, the rotation axis of the tiltrotor assembly 50 extends in the Z direction, with the tiltrotor assembly 50 in a vertical position.

Section 232 forms power unit of verting with the second of verting rotor subassembly 50, and power unit's focus is located the second and rolls over the axis of rotation of section 232 for power unit's focus can not be changed in verting rotor subassembly 50's rotation, can guarantee power unit's focus stability, and ensure the efficiency nature of the lift that hangs down. The center of gravity of the power mechanism is located on the rotation axis of the second turning section 232, so that the driving torque of the motor for driving the tilting connection member 30 (fig. 2) to rotate can be effectively reduced, the weight of the motor for driving the tilting connection member 30 to rotate is reduced, and the overall weight of the aircraft 1 is reduced.

In the present embodiment, the number of tilt rotor assemblies 50 is two, and each tilt rotor assembly 50 is correspondingly disposed at one turning portion 23, that is, two tilt rotor assemblies 50 are respectively connected to two corresponding second turning sections 232. The rotation direction of the tilt rotor blades 512 of the two tilt rotor assemblies 50 is opposite, so that the direction of the anti-torque generated by the rotation of the two tilt rotor blades 512 is opposite, and therefore the anti-torque is offset, the stability of the course of the aircraft 1 is improved, the deviation of the course can not occur, and the flight safety of the aircraft 1 is ensured.

In this embodiment, two tilt rotor assemblies 50 are symmetrically disposed about the center of gravity of aircraft 1 such that tilt rotor assemblies 50 do not change the position of the center of gravity of aircraft 1 during tilting, increasing the stability of aircraft 1 during flight.

With continued reference to fig. 1 and 2, the coaxial dual-rotor assembly 60 is rotatably disposed in the receiving cavity 11 and exposed at the opening 13. The coaxial dual-rotor assembly 60 can extend out of the fuselage 10 in the rotation process to provide lift force, and can be completely accommodated in the accommodating cavity 11 when the rotation is stopped, so that the resistance of the aircraft 1 in the flight process is reduced, and the high-speed flat flight of the aircraft 1 is facilitated. The coaxial dual-rotor assembly 60 is mainly responsible for helping the horizontal control surface to control the pitching balance of the whole machine and enhancing the stability of the whole machine.

The coaxial dual-rotor assembly 60 includes a mounting shaft 61, a first blade 63, and a second blade 65, wherein the mounting shaft 61 is connected to the fuselage 10, for example, the mounting shaft 61 may be connected to the fuselage 10 by a connecting rod located in the receiving cavity 11, and the first blade 63 and the second blade 65 are rotatably connected to opposite ends of the mounting shaft 61, respectively.

The first blade 63 and the second blade 65 are sequentially arranged in the third direction, that is, the first blade 63 and the second blade 65 are sequentially arranged in the Z direction to provide lift. In this embodiment, the rotation direction of the first blade 63 is opposite to the rotation direction of the second blade 65, so that the counter torque generated by the rotation of the first blade 63 and the counter torque generated by the rotation of the second blade 65 can be mutually offset, the heading stability of the aircraft 1 is improved, and the flight safety of the aircraft 1 is ensured. In this embodiment, the first blade 63 and the second blade 65 may be driven by two motors, respectively. In other embodiments, the first blade 63 and the second blade 65 may also be driven by one motor.

Two tiltrotor assemblies 50 constitute a quad-rotor module with co-axial dual-rotor assembly 60. In this embodiment, the center of gravity of aircraft 1 is located at the center of the quad-rotor module, that is, the distance between the center of coaxial dual-rotor assembly 60 and the center of gravity of aircraft 1 is equal to the distance between tilt-rotor assembly 50 and the center of gravity of aircraft 1, so that the position of the center of gravity of aircraft 1 is not affected by the state of coaxial dual-rotor assembly 60 and the state of tilt-rotor assembly 50, and the stability of aircraft 1 in the flight process is enhanced. The state of the coaxial dual-rotor assembly 60 includes an operating state and a stopped state, wherein the operating state refers to the rotation of the first blade 63 and the second blade 65 of the coaxial dual-rotor assembly 60; the stopped state refers to the first blade 63 and the second blade 65 of the coaxial dual rotor assembly 60 being fixed. Tiltrotor assembly 50 is in a state that includes a rotor state and a fixed-wing state.

In one embodiment, the center of coaxial dual rotor assembly 60 is spaced from the center of gravity of aircraft 1 by a greater distance than tilt rotor assembly 50 is spaced from the center of gravity of aircraft 1, i.e., coaxial dual rotor assembly 60 maintains a longer moment arm with the center of gravity, such that the overall primary source of lift for aircraft 1 is concentrated in tilt rotor assembly 50.

Referring to fig. 5 and 6, the aircraft 1 further includes a folding mechanism 70, the folding mechanism 70 is disposed on the fixing portion 21, and the folding portion 23 can be folded relative to the fixing portion 21 through the folding mechanism 70, so that the folding of the folding portion 23 relative to the fixing portion 21 is smoother, and the folding efficiency of the folding portion 23 is improved. The folding mechanism 70 may include a first rotating shaft 71 and a second rotating shaft 73, wherein the folding portion 23 is rotated along a first rotating axis by the first rotating shaft 71, and the folding portion 23 is rotated along a second rotating axis by the second rotating shaft 73, so as to fold the wing assembly 20 in steps, and reduce damage to the wing assembly 20 during the folding process. The folding mechanism 70 may rotate the folding portion 23 automatically or manually, for example, the first rotating shaft 71 and the second rotating shaft 73 may be driven by a motor, or the first rotating shaft 71 and the second rotating shaft 73 may be driven manually.

Referring to fig. 6 and 7, in the present embodiment, the aircraft 1 further includes a tail 90, the tail 90 is substantially "T" shaped, and the tail 90 is connected to the fuselage 10. The arrangement of the tail wings 90 can avoid wake flow interference, and the operation efficiency of the horizontal tail is improved.

The rear wing 90 includes a horizontal wing 92 and a vertical wing 94, wherein the horizontal wing 92 extends in the same direction as the wing assembly 20, i.e., the horizontal wing 92 extends in the X direction, and the vertical wing 94 is connected between the fuselage 10 and the horizontal wing 92 and extends upward, i.e., the vertical wing 94 extends in the Z direction, to control the pitch, yaw, and tilt of the aircraft 1, thereby controlling the attitude of the aircraft 1.

The aircraft 1 further comprises a wheel set 100, the wheel set 100 is arranged at the bottom of the fuselage 10, and the fuselage 10 can move along the ground through the wheel set 100, so that the aircraft 1 can run on the ground, and the transition requirement of the aircraft 1 is met. In this embodiment, the wheel set 100 can be driven by an in-wheel motor.

The working conditions of the aircraft 1 provided by the invention are explained below.

The aircraft 1 mainly comprises three working states, namely 1, vertical take-off and landing; 2. performing level flight cruising; 3. and (5) ground transition.

With continued reference to fig. 1 and 2, the vertical takeoff condition includes:

1. the aircraft 1 is located on the road surface and is moved to the take-off and landing platform by the wheel sets 100.

2. When aircraft 1 is located the platform of taking off and landing, through automatic or manual folding wing subassembly 20 and the rotor blade 512 that verts of expanding, set up and drive the second in the connecting piece 30 that verts of section 231 that verts and vert 90 upwards in the section 232 rotation of turning over of first section of turning over, thereby drive rotor component 50 that verts upwards and vert 90, make rotor component 50 that verts switch over to the rotor state, cover 17 is opened and is exposed coaxial two rotor components 60, control coaxial two rotor components 60 and begin work, namely, first paddle 63 and second paddle 65 rotate along opposite direction, the four rotor module begins to rotate, make aircraft 1 vertical lift.

The landing working conditions comprise:

1. when the aircraft 1 approaches a destination, the horizontal flying speed of the whole aircraft is reduced, the tilting connecting piece 30 drives the tilting rotor wing assembly 50 to tilt slowly from a horizontal position to a vertical position, the flap of the whole aircraft has a certain angle to improve the lift force of the wing assembly 20, the covering plate 17 is opened at the moment, the coaxial dual-rotor wing assembly 60 starts to work to keep the pitching balance of the whole aircraft, and the tilting rotor wing assembly 50 increases the upward lift force while reducing effective tension.

2. When aircraft 1 is located the airport top of taking off and land, drive and vert rotor subassembly 50 and upwards vert 90, four rotor modules begin to rotate, and aircraft 1 descends downwards.

Referring to fig. 2 and 4, the cruise conditions include:

1. after the aircraft 1 reaches a certain height, the tilting connecting piece 30 starts to slowly drive the tilting rotor assembly 50 to tilt to a horizontal position, the coaxial dual-rotor assembly 60 continuously works to keep the pitching balance of the whole aircraft, the flap of the whole aircraft has a certain angle to improve the lift force of the wing assembly 20, and the tilting rotor assembly 50 increases forward pulling force while providing effective lift force.

2. After the flying speed of aircraft 1 is higher than the flat flying speed, tilt rotor subassembly 50 rotates to horizontal position, aircraft 1 is in flat flying cruise operating condition this moment, coaxial dual rotor subassembly 60 stop work, first paddle 63 and second paddle 65 are fixed when rotatory to the position parallel with the Y direction, with the drag that flies that reduces, opening 13 is covered to cover plate 17, with the drag that flies that further reduces, the complete machine flap is withdrawed, aircraft 1's main lift source is changed wing subassembly 20 into by tilt rotor subassembly 50, tilt rotor subassembly 50 only provides forward pulling force this moment.

Referring to fig. 2 and 5, the ground transition condition includes:

1. when the aircraft 1 contacts the ground, the tilt rotor assemblies 50 rotate to the horizontal position, and the four rotor modules stop working;

2. folding of the wing assembly 20 and bending of the tiltrotor blades 512 is accomplished automatically or manually;

3. when the aircraft 1 leaves the take-off and landing platform, the wheelset 100 is used to complete the transition demand at short distances in the destination.

Wherein, folding of wing assembly 20 and the buckling of tiltrotor blade 512 include: first, tilt rotor blades 512 of tilt rotor assembly 50 are folded back in the Y direction; then, the folding mechanism 70 drives the folding portion 23 to rotate upward or downward 90 ° along the second rotation axis, and then rotate backward 90 ° along the first rotation axis; finally, a cover 17 at the rear end of the fuselage 10 covers the opening 13.

In summary, according to the aircraft 1 provided by the present invention, the coaxial dual-rotor assembly 60 is provided, and the rotation direction of the first blade 63 of the coaxial dual-rotor assembly 60 is opposite to the rotation direction of the second blade 65, so that the counter torque generated by the rotation of the first blade 63 and the counter torque generated by the rotation of the second blade 65 can be cancelled out, the stability of the heading of the aircraft 1 is improved, and the flight safety of the aircraft 1 is ensured.

Referring to fig. 8, the present invention further provides a coaxial dual-rotor assembly 60, wherein the coaxial dual-rotor assembly 60 includes a mounting shaft 61, a first blade 63 and a second blade 65, the first blade 63 and the second blade 65 are rotatably connected to opposite ends of the mounting shaft 61, respectively, and the rotation direction of the first blade 63 is opposite to the rotation direction of the second blade 65.

In summary, according to the coaxial dual-rotor assembly 60 provided by the present invention, the rotation direction of the first blade 63 is opposite to the rotation direction of the second blade 65, so that the counter torque generated by the rotation of the first blade 63 and the counter torque generated by the rotation of the second blade 65 can be cancelled out, and when the coaxial dual-rotor assembly 60 is used in an aircraft, the heading stability of the aircraft can be improved, and the flight safety of the aircraft can be ensured.

The above-mentioned embodiments only express several embodiments of the present invention, and the description thereof is more specific and detailed, but not construed as limiting the scope of the present invention. It should be noted that, for a person skilled in the art, several variations and modifications can be made without departing from the inventive concept, which falls within the scope of the present invention. Therefore, the protection scope of the present patent shall be subject to the appended claims.

Claims

1. An aircraft, characterized in that it comprises:

a body;

a wing assembly connected to the fuselage;

a tiltrotor assembly tiltably connected to the wing assembly; and

coaxial dual rotor subassembly, coaxial dual rotor subassembly is including installation axle, first paddle and second paddle, the installation hub connection in the fuselage, first paddle with the second paddle rotationally connect respectively in the relative both ends of installation axle, the rotation direction of first paddle with the rotation direction of second paddle is opposite.

2. The aircraft of claim 1, wherein the fuselage defines a cavity and an opening, the opening communicating with the cavity, the coaxial dual-rotor assembly being rotatably disposed within the cavity and exposed at the opening.

3. The aircraft of claim 2, wherein the fuselage further comprises a cover for covering the opening.

4. The aircraft of claim 3, wherein the fuselage further comprises a main body portion for positioning the wing assembly, the cover being rotatably positioned on the main body portion.

5. The aircraft of claim 1, wherein the wing assembly includes a fixed portion disposed on the fuselage and a flap portion movably connected to the fixed portion and pivotable relative to the fixed portion.

6. The aircraft of claim 5, wherein the fixed portion extends in a first direction, the deflection portion has a first axis of rotation and a second axis of rotation, the first axis of rotation extends in the first direction, the second axis of rotation extends in a second direction, the first and second paddles are arranged in series in a third direction, and the first, second, and third directions are perpendicular to one another.

7. The vehicle of claim 6, further comprising a folding mechanism disposed at the fixed portion, wherein the folding portion is foldable relative to the fixed portion by the folding mechanism.

8. The aircraft of claim 7, wherein the folding mechanism comprises a first pivot and a second pivot, the folding portion being rotated along the first pivot by the first pivot and the folding portion being rotated along the second pivot by the second pivot.

9. The vehicle of claim 5 wherein said turndown portion comprises a first turndown section and a second turndown section, said first turndown section being movably coupled to said fixed portion and being turndown relative to said fixed portion, said tiltrotor assembly being disposed in said second turndown section.

10. The aircraft of claim 9 wherein said tiltrotor assembly forms a power mechanism with said second folded section, said power mechanism having a center of gravity located at an axis of rotation of said second folded section.

11. The aircraft of claim 5, wherein the length of the fixed portion is 10% to 30% of the length of the wing assembly.

12. The aircraft of claim 1 further comprising a tilt connector rotatably disposed on said wing assembly, said tilt rotor assembly being drivingly connected to said tilt connector for changing the spatial angle upon actuation of said tilt connector.

13. The aircraft of claim 12 wherein the tilt rotor assembly comprises a tilt rotor body and a rotor mount, the rotor mount being in transmission connection with the tilt connection, the tilt rotor body being rotatably disposed in the rotor mount, the tilt rotor body comprising a tilt rotor blade, the tilt rotor blade being opposite the rotor mount being bendable.

14. The aircraft of any one of claims 1-13 wherein there are two of said wing assemblies, two of said wing assemblies being connected to opposite sides of said fuselage, one of said tiltrotor assemblies being provided for each of said wing assemblies, and two of said tiltrotor assemblies being symmetrically disposed about a center of gravity of said aircraft.

15. The aircraft of claim 14 wherein two of said tiltrotor assemblies and said coaxial dual-rotor assembly comprise a quad-rotor module, said aircraft having a center of gravity located at a center of said quad-rotor module.

16. The aircraft of claim 14 wherein a center of said coaxial dual rotor assembly is spaced from a center of gravity of said aircraft by a distance greater than a distance between a center of gravity of said tiltrotor assembly and said aircraft.

17. The aircraft of any one of claims 1-16, further comprising an empennage attached to the fuselage.

18. The aircraft of claim 17 wherein the tail wing includes a horizontal wing and a vertical wing connected between the fuselage and the horizontal wing and extending upwardly, the horizontal wing extending in the same direction as the wing assembly.

19. The vehicle according to any one of claims 1 to 18, further comprising a set of wheels arranged at the bottom of the fuselage, the fuselage being movable along the ground by means of said set of wheels.

20. A coaxial dual-rotor assembly, comprising an installation shaft, a first blade and a second blade, wherein the first blade and the second blade are rotatably connected to opposite ends of the installation shaft, respectively, and the rotation direction of the first blade is opposite to the rotation direction of the second blade.