WO2022037450A1 - 一种结合多轴旋翼的自旋翼飞行器 - Google Patents

一种结合多轴旋翼的自旋翼飞行器 Download PDF

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Publication number
WO2022037450A1
WO2022037450A1 PCT/CN2021/111949 CN2021111949W WO2022037450A1 WO 2022037450 A1 WO2022037450 A1 WO 2022037450A1 CN 2021111949 W CN2021111949 W CN 2021111949W WO 2022037450 A1 WO2022037450 A1 WO 2022037450A1
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WIPO (PCT)
Prior art keywords
rotor
aircraft
fuselage
axis
supply device
Prior art date
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PCT/CN2021/111949
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English (en)
French (fr)
Inventor
吴斌
王辉
Original Assignee
加拿大轻型航空有限公司
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from CN202021772816.3U external-priority patent/CN212861854U/zh
Priority claimed from CN202023122491.7U external-priority patent/CN214138980U/zh
Priority claimed from CN202011533132.2A external-priority patent/CN112722289A/zh
Application filed by 加拿大轻型航空有限公司 filed Critical 加拿大轻型航空有限公司
Publication of WO2022037450A1 publication Critical patent/WO2022037450A1/zh

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/40—Arrangements for mounting power plants in aircraft

Definitions

  • the invention belongs to the technical field of aircraft, and in particular relates to a self-rotating aircraft combined with a multi-axis rotor.
  • the autorotor is a type of rotorcraft that uses the autorotor as the lifting surface, the propeller push/pull or other energy supply methods as the forward power.
  • the rotor of the autogyro is driven by the forward flow during flight to achieve self-rotation to provide lift. Once its engine is stopped in the air, it can rely on the autorotor to land safely, so it has historically been called “the safest flying machine”.
  • the auto-rotor needs a certain taxiway to take off, it cannot take off and land vertically. When flying at a low altitude, it cannot lift up quickly if it encounters obstacles such as trees and telephone poles, and the auto-rotor can only keep one kind of forward movement.
  • the flight attitude is also unable to hover in the air and fly left and right, which leads to a significant reduction in the scope of application of this type of autogyro, and the inability to lift or quickly change the moving direction during use leads to many safety hazards.
  • the common multi-axis rotorcraft can solve the problem of vertical take-off and landing in a narrow space, but due to the design of its fixed bracket structure, it is easily affected by the airflow during vertical lift, and the resistance is large, resulting in unstable vertical lift and energy Large consumption.
  • the battery compartment of this type of aircraft is arranged in the fuselage, which is not conducive to the heat dissipation and replacement of the high-power battery pack, and the use range is narrow.
  • the use range is narrow.
  • can not flexibly deal with various use scenarios and other problems; further optimizing the multi-axis rotor auto-rotor aircraft to make it have higher stability, wider range of use, stronger endurance, etc. are also problems that those skilled in the art need to solve urgently.
  • the present invention provides a self-rotating aircraft combined with a multi-axis rotor.
  • a gyrocopter incorporating a multi-rotor comprising
  • the rotor is arranged on the top of the fuselage
  • a propulsion device which is arranged at the tail of the fuselage, and includes a propulsion paddle and a propulsion drive device that drives the propulsion paddle to rotate;
  • At least one group of multi-axis rotor systems is used to provide vertical lift for the fuselage; each group of the multi-axis rotor systems includes two rotor modules symmetrically arranged on both sides of the fuselage; each of the rotor modules It includes at least one rotor, each of which is correspondingly connected with a rotor drive device, and the rotor module is mounted on the fuselage through a bracket;
  • the landing gear is arranged on the bottom of the fuselage or on the bracket.
  • a flight power system which includes: a power control module, a fuel supply device, and an electric power supply device;
  • the power control module is respectively connected with the fuel supply device, the propulsion drive device, the rotor drive device and the power supply device;
  • the power supply device is controlled to provide power to the rotor drive device to drive the rotor to rotate; during parallel flight, the fuel supply device is controlled to provide power to the propulsion drive device, Or control the electric power supply device to provide electric power to the propulsion driving device, so as to drive the propeller to rotate.
  • the multi-axis rotor system is detachably installed on the fuselage; wherein, when the multi-axis rotor system is installed, a multi-axis rotor and spin-rotor combined aircraft is formed; when the multi-axis rotor system is disassembled , which constitutes a self-rotating aircraft.
  • each of the rotor modules includes two rotors, and the two rotors are arranged coaxially up and down.
  • the propulsion driving device is a fuel engine
  • a generator is provided in the fuselage, and the generator is connected to the fuel engine for charging the power supply device.
  • the bracket adopts a foldable and/or retractable bracket.
  • a battery compartment reinforcing rib is provided on the support frame at a position corresponding to the battery compartment, and the battery compartment is hung on the battery compartment reinforcing rib by a disassembly assembly; the disassembly assembly includes:
  • the radial insert is inserted into the connecting seat and the connecting head along the radial direction of the connecting seat, so as to realize the axial limit of the connecting head;
  • the power supply device discharges at high power and releases heat at the same time.
  • the installation of the power supply device inside the fuselage is not conducive to heat dissipation, so it is a power supply device.
  • the arrangement of the liquid cooling system will increase the weight and the complexity of the system; the present invention uses the air cooling to improve the heat dissipation capacity of the power supply device by hanging the power supply device externally, and the power supply device is directly exposed to the air.
  • the power supply device or the battery module in the battery compartment can be taken out and stored indoors to avoid serious power loss when the ambient temperature is very low, and to improve the battery life, so that there is no need for additional cooling system and thermal insulation system inside the fuselage.
  • FIG. 2 is a top view of a spin-rotor aircraft combined with a multi-axis rotor in Embodiment 1 of the present invention
  • FIG. 3 is a schematic diagram of the battery compartment installed on the bracket in Embodiment 1 of the present invention.
  • Embodiment 4 is a schematic top view of the battery compartment in Embodiment 1 of the present invention.
  • FIG. 5 is a schematic structural diagram of a disassembled assembly in Embodiment 1 of the present invention.
  • FIG. 6 is a schematic diagram of a spin-rotor aircraft combined with a multi-axis rotor in Embodiment 2 of the present invention
  • FIG. 8 is a schematic diagram of a spin-rotor aircraft combined with a multi-axis rotor in Embodiment 4 of the present invention.
  • FIG. 9 is a schematic diagram of a self-rotating aircraft combined with a multi-axis rotor in Embodiment 4 of the present invention.
  • FIG. 10 is a top view of a spin-rotor aircraft incorporating a multi-axis rotor in Embodiment 4 of the present invention.
  • FIG. 11 is a schematic diagram of a spin-rotor aircraft incorporating a multi-axis rotor in Embodiment 5 of the present invention.
  • FIG. 12 is a schematic diagram of a spin-rotor aircraft incorporating a multi-axis rotor in Embodiment 6 of the present invention.
  • FIG. 13 is a schematic diagram of a spin-rotor aircraft incorporating a multi-axis rotor in Embodiment 7 of the present invention.
  • FIG. 14 is a schematic diagram of a spin-rotor aircraft incorporating a multi-axis rotor in Embodiment 8 of the present invention.
  • FIG. 15 is a schematic diagram of a spin-rotor aircraft incorporating a multi-axis rotor in Embodiment 9 of the present invention.
  • FIG. 16 is a schematic diagram of a spin-rotor aircraft combined with a multi-axis rotor according to Embodiment 10 of the present invention.
  • FIG. 17 is a schematic diagram of a spin-rotor aircraft incorporating a multi-axis rotor in Embodiment 11 of the present invention.
  • FIG. 18 is a schematic diagram of the vertical flight principle of a spin-rotor aircraft combined with a multi-axis rotor in Embodiment 1 of the present invention
  • FIG. 19 is a schematic diagram of the flight principle of the spin-rotor aircraft combined with multi-axis rotors in different attitudes according to Embodiment 1 of the present invention.
  • the present invention provides a gyrocopter incorporating a multi-axis rotor, comprising:
  • the system is used to provide vertical lift for the fuselage 4; each group of multi-axis rotor systems includes two rotor modules 1 symmetrically arranged on both sides of the fuselage 4; each rotor module 1 includes at least one rotor 11, each rotor 11
  • a rotor drive device 13 is connected, and the rotor module 1 is installed on the fuselage 4 through a bracket; the landing gear 6 is arranged on the bottom of the fuselage 4 or on the bracket.
  • the landing gear 6 can adopt the front three-point non-retractable landing gear, the front wheel can be controlled by pedals to realize ground turning, and the main wheel is provided with synchronous brake.
  • the rotor in this embodiment is arranged on the top of the fuselage, and is inclined backward by about 4°, so that the rotor blades of the rotor are always at about 4° with the windward angle. It is further explained that if the rotor rotates counterclockwise, the blades on the right side of the fuselage always have the leading edge forward due to the inclination of 4°. Since the blades have a certain airfoil, the resistance generated is small; the blades on the left side of the fuselage always With the trailing edge forward, the blade as a whole generates a larger angle of attack with the incoming flow, which can generate a larger lift force.
  • the propeller is arranged at the tail of the fuselage to provide power for the aircraft to move forward in level flight.
  • the multi-axis rotor system in this embodiment includes rotor modules 1 symmetrically arranged on both sides of the fuselage, and each rotor module 1 has at least one rotor 11 .
  • the flight attitude of the aircraft is changed by changing the rotational speed of each rotor 11 .
  • the multi-axis rotor system in this embodiment is detachably installed on the fuselage 4; wherein, when the multi-axis rotor system is installed, it constitutes a multi-axis rotor and spin-rotor combined aircraft; when the multi-axis rotor system is disassembled, it constitutes a self-rotor aircraft.
  • this embodiment uses the reconfigurable design concept and according to the design requirements of the aircraft, the pre-determined aircraft components are refitted into aircraft with required functions, that is, the redesigned reconfigurable light aircraft in this embodiment,
  • the vertical take-off and landing system is composed of multi-rotor rotor systems of different structures, and it becomes a hybrid light aircraft capable of both vertical take-off and landing and runway take-off and landing.
  • the reconfigurable aircraft in this embodiment is applied to a manned aircraft, so that when a multi-axis rotor system is installed, a manned aircraft combined with multi-axis rotors and rotors is formed. When the multi-axis rotor system is disassembled, a self-rotating manned aircraft is formed.
  • the tail portion of the fuselage in this embodiment may be provided with a tail fin 15, or a tail-less fuselage may be directly used.
  • the tail fin 15 may be a flat tail, a V-shaped tail fin, or a tail with other configurations, which is not limited here.
  • the fuselage is a tailless fuselage, multiple sets of multi-axis rotor systems are arranged on the tailless fuselage.
  • the top of the fuselage 4 is provided with a rotor 5, and the bottom is provided with a landing gear 6.
  • the propeller 3 in this embodiment is not limited to being provided at the rear end of the fuselage 4. In other embodiments, it can also be provided on the head of the fuselage 4.
  • Two multi-axis rotor systems are arranged symmetrically about the fuselage 4 longitudinally.
  • the multi-axis rotor system in this embodiment is detachably mounted on the fuselage 4, which is convenient for retrofitting into a desired aircraft configuration.
  • a flight power system which at least includes a power control module, a fuel supply device, and an electric power supply device 7 .
  • the power control module is respectively connected with the fuel supply device, the propulsion drive device 31, the rotor drive device 13 and the power supply device 7; according to the selected flight state, the power supply device 7 is controlled to provide electric drive to the rotor drive device 13 during vertical take-off and landing The rotor 11 rotates; during parallel flight, the fuel supply device is controlled to provide power to the propulsion drive device 31 to drive the propulsion paddle 3 to rotate, or the power supply device 7 is controlled to provide power to the propulsion drive device 31 to drive the propulsion paddle 3 to rotate.
  • the flight power system in this embodiment includes at least one set of power supply devices 7 and is arranged symmetrically with respect to the average center of gravity of the entire aircraft.
  • the power supply device 7 in this embodiment can be built inside the body 4 , or mounted on a bracket or outside the body 4 .
  • the power supply device 7 is mounted outside the bracket or the fuselage 4 .
  • the power control module, the fuel supply device, and the propulsion drive device 31 in this embodiment are arranged inside the fuselage 4 .
  • the bracket includes a rotor frame 12 and a support frame 9 , and the rotor module 1 is mounted on the support frame 9 through the rotor frame 12 .
  • the power supply device 7 is mounted on the outside of the support frame 9 or the fuselage 4 .
  • the power supply device 7 is hung below the support frame.
  • the power supply device 7 is creatively suspended outside the aircraft.
  • the advantages are: 1. There is no need to install another power supply device in the fuselage to occupy the cramped space inside the fuselage, and there is no need to lengthen the fuselage. Expand the internal space of the aircraft to avoid increasing the design difficulty and production cost of the aircraft; 2.
  • the power supply device is externally attached to the aircraft, which is conducive to quick disassembly and replacement. After the aircraft reaches the destination, the backup power supply device can be directly replaced without waiting for charging, which effectively improves work efficiency. Efficiency; 3. When the rotor drive device of the rotor is working, the power supply device is used for high-power discharge and high-power heat dissipation.
  • the power supply device When the power supply device is installed inside the fuselage, it is not conducive to the heat dissipation of the battery, and the liquid cooling system for the power supply device will increase. weight and increase the complexity of the system; in the present invention, the power supply device is directly exposed to the air by externally hanging the power supply device, and the cooling capacity of the power supply device is improved by using the cold air outside the aircraft. When the aircraft is not sailing, the power supply device can be taken out. It is stored indoors to avoid serious power loss in the power supply device when the ambient temperature is very low, and to improve the service life, so there is no need for additional cooling system and thermal insulation system inside the fuselage.
  • the multi-axis rotor system in this embodiment includes at least two rotor modules 1 , that is, there are two rotor modules 1 on each side of the fuselage 4 , so that when adjusting the flying attitude of the aircraft, it can be adjusted from four degrees of freedom. Control and enhance the flight stability and flexibility of the aircraft.
  • the number of groups set by the rotor module 1 is not limited to the above, and the rotor module 1 can be adjusted according to the configuration of the aircraft, the required lift of vertical take-off and landing, the stability sensitivity of regulation, etc. number.
  • the multi-axis rotor system includes four rotor modules 1 or five rotor modules 1 or the like.
  • the rotor modules 1 located on the same side of the fuselage 4 share a bracket.
  • the rotor modules 1 located on the same side of the fuselage 4 share a rotor frame 12, and the two rotor modules 1 are respectively arranged at both ends of the rotor frame 12 and are located on the same horizontal plane.
  • the average center of gravity of the fuselage 4 is symmetrically arranged, and the corresponding two rotor modules 1 in the two multi-axis rotor systems are symmetrical to each other in the longitudinal direction of the fuselage, so that the rotor modules 1 on both sides of the fuselage 4 and the front and rear directions of the fuselage 4 are symmetrical to each other,
  • the pulling force is basically the same, and the force of the aircraft is balanced, so that there is no need to consume too much energy to adjust the balance between the rotors 11, which greatly reduces the energy loss.
  • the relative positions of the two rotor modules 1 in each multi-axis rotor system are not limited to the above, and corresponding layout adjustments can be made according to the structure of the fuselage and practical applications.
  • each rotor module may also correspond to a rotor frame.
  • each rotor module 1 includes a group of rotors, so that the entire aircraft includes four rotor modules, and further includes four groups of rotors, and each rotor is connected to a rotor drive device
  • the rotor drive devices are A, B, C, and D respectively, when the aircraft moves vertically, the rotors of the adjacent two groups rotate in opposite directions, and the two rotor drive devices on the diagonal The driving rotors rotate in the same direction and at the same speed.
  • one group of the rotor drive devices on the diagonal needs to drive the rotor to rotate clockwise in order to generate upward lift; the other group of the rotor drive devices on the diagonal needs to The rotor is driven in a counterclockwise direction to generate upward lift. It can also be understood that when each group of rotors generates lift, the airflow is blown downward to provide upward pulling force.
  • each rotor control can adopt the above-mentioned flight attitude control principle. This embodiment will not further elaborate on this.
  • the preferred rotor module 1 includes two rotors 11, and the two rotors 11 are arranged coaxially up and down to form a coaxial double rotor, which greatly improves the performance of each propeller module.
  • its vertical take-off and landing lift becomes stronger and the structure is more compact, the space occupied by it is smaller, and the weight efficiency is higher;
  • the coaxial dual rotor method makes the space between the two coaxial rotors in the hovering state. The aerodynamic disturbance will have a beneficial effect and improve the hovering efficiency.
  • the load-bearing beams inside the support frame 9 are bolted to the beams inside the fuselage, and the connection between the support frame 9 and the fuselage 4 can be completed in the nacelle, which is convenient for disassembly and maintenance.
  • a bolt connection combined with a bolt connection can also be used; of course, joints can be provided on the fuselage, and the support frame is connected to the fuselage through the joints, which is not limited here.
  • one end of the support frame 9 is connected to the fuselage 4, and the other end is tilted forward along the longitudinal direction of the fuselage to the head of the fuselage, forming a forward-swept layout, so that the support frame 9 can avoid
  • the position of opening the aircraft cabin door is arranged so that the two rotor modules located on the same side of the fuselage can be symmetrical about the average center of gravity of the fuselage, and the power supply device 7 mounted under the support frame 9 is as close as possible to the average center of gravity of the fuselage. Longitudinal, the weight distribution of the whole machine is more uniform, which greatly improves the stability and balance of the aircraft.
  • the setting direction and position of the support frame 9 are not limited to those described above or shown in the figures. Corresponding adjustments need to be made according to the actual layout of the fuselage to prevent the support frame 9 from hindering the entry and exit of the aircraft and ensure the rotor module.
  • the layout of the power supply device 7 is reasonable.
  • the support frame 9 may also be a circle-shaped frame symmetrically arranged along the fuselage or two brackets inclined toward each other in a figure-eight shape.
  • the support frame 9 in this embodiment includes two power supply devices 7 , and the two power supply devices 7 are arranged along the length direction of the support frame 9 .
  • the two power supply devices 7 are arranged along the extension direction of the support frame 9, so that the power supply device 7 is as close as possible to the longitudinal direction of the average center of gravity of the fuselage, and the weight distribution of the whole machine is more uniform. Greatly improved the stability and balance of the aircraft.
  • the layout of the power supply device 7 is not limited to the above.
  • the battery compartment reinforcement rib is provided on the support frame 9 at the position corresponding to the battery compartment 8, and the battery compartment 9 is hung on the battery compartment reinforcement rib through the dismantling assembly 2, so as to prevent the support frame 9 from being overstressed locally, thereby extending the length of the battery compartment. The service life of the support frame 9.
  • the disassembly assembly 2 in this embodiment includes: a connecting base 21 having a socket, and the connecting base 21 is fixed on the support frame 9 ; a connecting head 23 is arranged on the battery compartment 8 , and the connecting head 23 is connected to
  • the sockets are matched, and the connector 23 is inserted into the socket to realize the radial limit of the connector; the radial insert 24 is inserted into the connector seat 21 and the connector 23 along the radial direction of the connector seat 21 to realize the axial limit of the connector.
  • Position; the fixing member 25 is used to limit the radial movement of the radial insert 24 from the connecting seat 21 and the connecting head 23 .
  • the quick disassembly and replacement of the battery compartment 8 and the power supply device 7 can be realized by the disassembly structure.
  • the specific structure of the disassembly assembly is not limited to the above or shown in the drawings, for example, other connection assemblies may also be used.
  • the dismantling assembly 2 further includes a reinforcing frame 22 , and the battery compartment 8 is surrounded by a reinforcing frame 22 ; To improve the strength of the battery compartment 8.
  • the radial inserts 24 and the fixing pieces 25 are all plug-in structures.
  • the specific structures of the radial inserts and fixing pieces are not limited to the above, for example, other connecting pieces or Nuts, bolts, etc.
  • the cross section of the preferred support frame 9 is an elliptical streamlined section, the bottom of which can provide a larger area for suspending the power supply device 7, and the streamlined shape of the elliptical streamlined section is a rounded front and a rear tip, Slightly shaped like a water droplet, the support frame 9 supports the power supply device 7 more stably and with higher strength; and the streamlined support frame 9 can generate lift to supplement the lift of the aircraft without increasing air resistance.
  • the cross-sectional shape of the support frame 9 is not limited to the above or shown in the figures, for example, the cross-section of the support frame 9 may also be circular.
  • the flight power system in the autorotor combined with the multi-rotor rotor includes a power control module, a power supply device 7 and a fuel power device.
  • the power control module can control the power supply device 7 and the fuel power device to output power together.
  • the electric power supply device 7 and the fuel power device can both provide power for the aircraft, and the power of the two can be jointly output or independently output to the corresponding device.
  • the power supply device 7 provides energy for the rotor drive device 13, so that the rotor drive device 13 drives the rotor to rotate to realize vertical take-off and landing of the aircraft, low-speed flight in multi-axis mode, etc.
  • the fuel engine of the fuel power device provides power for the propulsion drive device 31.
  • the electric power supply device 7 and the fuel power device can also jointly provide power for the rotor drive device 13 and the propulsion drive device 31 .
  • the power of the autorotor combined with the multi-rotor is all derived from the electric power system, and the control system controls the electric power device to provide power for the rotor drive device 13 and the propulsion drive device 31 , wherein the engine of the propulsion drive device 31 provides power. for the electric engine.
  • the power of the electric power system comes from the above-mentioned electric power supply device 7 and/or an additionally provided battery pack, and there is no need to provide a fuel power device.
  • the rotor drive device 13 equipped with each rotor 11 in this embodiment is a drive motor, and the drive motor is arranged on a bracket (which may be on a support frame or a rotor frame).
  • the drive motor can also be set in other positions, there is no restriction here, and the best setting position can be selected according to the actual situation.
  • the propulsion drive device 31 in this embodiment can be a fuel engine or an electric engine.
  • the fuel supply device provides power
  • the propulsion drive device 31 is an electric engine
  • the power supply device 7 provides power.
  • the propulsion driving device is a fuel engine
  • a generator is provided in the fuselage, and the generator is connected to the fuel engine for charging the power supply device 7.
  • the power supply device 7 can be charged to ensure that the Energy supply during vertical lift. It is worth noting that when the propulsion drive device 31 is a fuel engine, the engine needs to be preheated to start. When the propulsion drive device 31 is preheated, the propeller 3 rotates to generate a certain thrust to make the aircraft slide forward.
  • the braking device needs to be used. Hold the fuselage so that the fuselage does not slide forward when it is on the ground; when the aircraft starts to take off away from the ground, the rotation speeds of the two sets of multi-axis rotor systems in the front and rear of the fuselage 4 are inconsistent, and then a certain backward force is applied to the fuselage 4 It is offset with the thrust generated by the propeller 3 to ensure that the aircraft can take off vertically. Therefore, when the propulsion driving device 31 is a fuel engine, at least two sets of multi-shaft rotor systems need to be provided.
  • the propulsion drive device 31 is an electric engine, preheating is not required, and the propeller 3 can be started after vertical take-off, so as not to generate forward thrust to the fuselage 4, and at least one set of multi-axis rotor systems only needs to be installed at this time.
  • the preferred bracket (which may be a support frame, a rotor frame, or a support frame and a rotor frame) is a foldable frame. When the aircraft is not in operation, the frame is folded and placed to save its storage space.
  • the preferred bracket (which can be a support frame, a rotor frame or a support frame and a rotor frame) is a retractable frame, so that the length of the frame can be adjusted to obtain the best flying effect; or when the aircraft is not in operation,
  • the retractable stand saves its storage space.
  • the direct support can also be a foldable and retractable support, which is not limited here.
  • the autorotor aircraft combined with multi-axis rotors can achieve various flight attitudes.
  • the following example illustrates the process of the aircraft achieving various flight attitudes.
  • the rotor drive devices 13 of the plurality of rotors 11 are activated at the same time, and the multiple rotors 11 rotate at the same time and generate enough lift, so that the aircraft can take off vertically; it should be noted that the propulsion drive device 31 During warm-up, the rotation of propeller 3 generates a certain thrust to make the fuselage slide forward.
  • the braking device cannot provide resistance to prevent the fuselage from moving forward. At this time, adjust the front and rear groups of the fuselage.
  • the rotation speed of the rotor 11 makes the rotation speed inconsistent, and then the certain backward force applied to the fuselage is offset by the thrust generated by the propeller 3, thereby ensuring that the aircraft can take off vertically.
  • the forward thrust of the propulsion paddle 3 increases, which pushes the aircraft to fly forward. Accelerate, this lift increases rapidly, the multiple rotors 11 reduce the rotational speed at the appropriate time, and the lift provided gradually decreases accordingly.
  • the lift generated by the 5 blades of the rotors is sufficient to support the horizontal flight of the aircraft.
  • the aircraft The horizontal flight at a stable altitude is maintained entirely by the lift generated by the rotor 5 .
  • the rotational speed of the rotor 11 is increased while the rotational speed of the propulsion drive device 31 is decreased, until the lift generated by the rotor 11 is equal to the weight of the aircraft, the rotational speed of the rotor 11 is kept unchanged, and the aircraft remains in the hovering state.
  • the aircraft can also fly left and right by adjusting the rotational speed of the rotors 11 on both sides of the fuselage.
  • the preferred aircraft is a manned aircraft.
  • the aircraft constitutes a manned aircraft combined with multi-axis rotors and rotors; of course, in other embodiments, After dismantling the multi-rotor system, the aircraft constitutes a self-rotating vehicle, and it is worth noting that the auto-rotor can still carry people.
  • the appropriate configuration can be selected according to the needs of use to expand the application scope of the aircraft. No matter whether the multi-axis rotor system is installed or not, it is necessary to ensure that the constituted aircraft can carry people and meet the airworthiness certification standards for manned aircraft.
  • the rotors of different configurations can be replaced according to the maximum allowable take-off weight of the aircraft. For example, for aircraft with a lighter maximum take-off weight, a smaller rotor needs to be replaced.
  • Adapt to the corresponding rotor according to the application change the model flexibly, and use it in a wider range. It should be noted that the different configurations of the rotors are matched with the interface of the fuselage to ensure that the different configurations of the rotors can be replaced on the same fuselage.
  • Embodiment 1 is an improvement based on Embodiment 1. Referring to FIG. 6 , the difference between this embodiment and Embodiment 1 is:
  • the bracket further includes a support beam 10, one end of the support beam 10 is connected with the support frame 9, and one end is connected with the fuselage 4, and the support beam 10 is connected with the fuselage 4 and the support frame 9 in a triangular structure, which effectively strengthens the The strength of the bracket.
  • one end of the support beam is connected with the load-bearing beam of the rotor inside the fuselage, and one end is connected with the support frame 9 in a triangular structure.
  • the support beam can also be The connection with the main force transmission beam of the fuselage is in a triangular structure, that is, the support beam 10 and the support frame 9 are located on the same horizontal plane.
  • This embodiment is an improvement based on Embodiment 2.
  • the difference between this embodiment and Embodiment 1 is that in this embodiment, the cross section of the support frame 9 is circular.
  • the support with an elliptical streamline cross-section will also generate lift, which will have a certain impact on the balance and stability of the whole aircraft. Therefore, in this embodiment, by setting the cylindrical support frame 9, the support frame 9 is affected by The area of the wind becomes smaller, so that the lift force received by the support frame 9 during flight becomes smaller or even disappears, which reduces the influence on the balance and stability of the aircraft, and is conducive to more accurate flight control of the aircraft.
  • Other structures and their connection relationships in this embodiment are the same as those in Embodiment 1, and will not be repeated here.
  • This embodiment is an improvement made on the basis of Embodiment 1.
  • the difference between this embodiment and Embodiment 1 is:
  • the multi-axis rotor system includes four rotor modules 1 , two rotor modules 1 share a rotor frame 12 and a support frame 9 , wherein the two rotor modules 1 are arranged on the fuselage head , the other two rotor modules 1 are arranged at the average center of gravity of the fuselage 4, and the length and width of the rotors in the two rotor modules 1 at the head of the fuselage 4 are smaller than the rotors 11 in the other two rotor modules 1, thereby further improving the aircraft.
  • it can effectively reduce the influence of the rotor module at the head of the fuselage on the average center of gravity of the fuselage.
  • the four rotor modules 1 located on the same side of the fuselage share one support frame 9 in two pairs, and the four sets of rotor modules 1 are respectively arranged on the head and tail of the fuselage 4 .
  • the flight attitude of the aircraft is regulated from four degrees of freedom to enhance the stability and flexibility of the aircraft's flight.
  • the rotor module 1 includes one rotor 11 .
  • the number of rotors 11 included in the rotor module 1 is not limited to the above or shown in the figures.
  • the rotor module 1 may also include two rotors 11 .
  • the rotor drive device 13 equipped with each rotor 11 is a drive motor, and the drive motor is built into a bracket (which may be a support frame or a rotor frame).
  • the drive motor can also be set in other positions, there is no restriction here, and the best setting position can be selected according to the actual situation.
  • the spin-rotor aircraft combined with the multi-axis rotor provided by this embodiment can realize various flight attitudes such as vertical take-off and landing, hovering in the air, air cruising, and side-to-side flying through the cooperation of the multi-axis rotor system, the rotor and the propulsion device;
  • the autorotor combined with the multi-rotor provided by this embodiment can take off without a runway, and can more flexibly select an appropriate flight attitude during operation, thereby greatly improving the application range and safety of the autorotor.
  • This embodiment is an improvement made on the basis of Embodiment 1.
  • the difference between this embodiment and Embodiment 1 is:
  • the rotor module 1 further includes a protective cover 14, and the protective cover 14 is arranged on the peripheral side of the rotor, so as to protect the rotor 11 without affecting the vertical airflow of the rotor, It can prevent people or objects from being accidentally injured when the rotor 11 rotates, and of course, it can also prevent the rotor 11 from loosening and falling off and hitting the fuselage when the rotor 11 is rotated at a high speed.
  • each rotor 11 is equipped with a protective cover 14 , and the cross section of the protective cover 14 is circular and is arranged around the circumference of the rotor 11 , wherein the protective cover 14 is arranged on the rotor frame 12 .
  • the specific shape of the protective cover 14 is not limited to those described above or shown in the figures.
  • the cross section of the protective cover 14 can also be an oval protective cover.
  • Embodiment 1 is an improvement based on Embodiment 1. Referring to FIG. 12 , the difference between this embodiment and Embodiment 1 is:
  • the support frame 9 is a wing structure.
  • the wing structure When the aircraft is taxiing, taking off and landing, the wing structure further provides lift for the aircraft, so that the power requirement for the rotor is lower.
  • the specific structure of the bracket is not limited to the above or shown in the drawings.
  • the multi-axis rotor system includes four rotor modules 1, two rotor modules 1 share a rotor frame 12, and the two rotor modules 1 are symmetrically arranged with respect to the average center of gravity of the fuselage, which greatly improves the aircraft's performance. Stability and balance.
  • each rotor module 1 includes two rotors 11, and the two rotors 11 are arranged coaxially up and down to form a coaxial double rotor.
  • the coaxial dual-rotor approach makes each rotor module 1 larger in total power, small in size, compact in structure and high in weight efficiency.
  • two sets of multi-axis rotor systems are provided, each rotor module 1 is mounted on the fuselage 4 through a bracket (including the support frame 9 and the rotor frame 12 ), the rotor module 1 includes two rotors 11, and The two rotors 11 are coaxially arranged up and down on the same bracket.
  • there is no limitation on the number of sets of multi-shaft rotor systems for example, three sets of multi-shaft rotor systems may also be set.
  • each rotor module 1 in Embodiment 4 includes one rotor 11 and is provided with four sets of rotor modules 1, the rotors 11 located on the same side of the fuselage 4 share one support frame 9 in pairs.
  • the coaxial double rotor is used to greatly increase the total power of each rotor module 1.
  • the vertical take-off and landing lift becomes stronger and the structure is more compact, and the space occupied by it is almost doubled compared to Example 4.
  • the weight efficiency is higher; further, the coaxial dual rotor method makes the aerodynamic interference between the two coaxial rotors have a favorable impact in the hovering state, and improves the hovering efficiency.
  • this embodiment is an improvement based on Embodiments 1 and 4.
  • the differences between this embodiment and Embodiments 1 and 4 are:
  • This embodiment is a modification made on the basis of Embodiment 4.
  • the difference between this embodiment and Embodiment 4 is:
  • the rotor modules 1 are symmetrically distributed on both sides of the fuselage 4 in a figure-eight shape.
  • four sets of multi-axis rotor systems are provided, and the four rotor modules 1 located on both sides of the fuselage are arranged in a figure-eight shape along the axial direction of the fuselage.
  • the multi-axis rotor system is arranged in a figure-eight shape, thereby increasing the distance between the rotor 11 at the tail of the fuselage and the propeller 3, so that when the two rotate at the same time, the airflow generated around the two will not disturb or interfere with each other. smaller, which is beneficial to improve aerodynamic efficiency.
  • the multiple groups of rotors arranged in a figure-eight shape are connected by two support frames, and four landing gears are symmetrically arranged on the lower sides of the two supports.
  • the rotor modules located on the same side of the fuselage share a support frame, and the support frames on both sides are arranged in a figure-eight shape.
  • four landing gears are arranged facing each other on the undersides of the two brackets. The four landing gears are arranged under the two brackets in the figure-eight arrangement, forming a rectangular-like arrangement, which provides a more balanced and stable support force when the aircraft lands, so that the aircraft can land more smoothly and improve user comfort.
  • This embodiment is a modification made on the basis of Embodiment 4.
  • the difference between this embodiment and Embodiment 4 is:
  • four groups of multi-axis rotor systems are provided, wherein two groups of multi-axis rotor systems are circumferentially arranged around the fuselage 1, and four rotor modules are connected by a ring bracket to form an annular portion; the other two groups
  • the multi-axis rotor system is arranged outside the annular portion, and the two rotor modules located on the same side are connected to both sides of the annular portion through an arc-shaped bracket.
  • the multi-axis rotor system is arranged on the top of the fuselage.
  • the location where the multi-axis rotor system is arranged is not limited.
  • the multi-rotor system can also be arranged in the middle of the fuselage or at the bottom of the fuselage.
  • this embodiment has an advantage in that the arrangement of the screw components in this embodiment is beneficial to improve the stable shape of the aircraft, and it is convenient to control the rotational speed of each rotor to control the flight state of the aircraft.
  • This embodiment is a modification made on the basis of Embodiment 1. The difference between this embodiment and Embodiment 1 is:
  • the fuselage is a tailless fuselage, and at least two sets of multi-axis rotor systems are arranged on the tailless fuselage.
  • the fuselage is provided with the tail wing 11
  • the aircraft is controlled by the tail wing to complete the movements such as yaw or pitch during cruise forward flight, the rotor is in a closed state at this time.
  • the tail is omitted, at least two sets of multi-axis rotor systems must be continuously turned on when the aircraft is flying, and the flight attitude of the aircraft can be controlled by adjusting the speed difference of multiple rotors.
  • the tail wing structure is omitted in the fuselage without tail wing in this embodiment.
  • the flight attitude of the aircraft such as yaw and pitch can be changed by adjusting the rotational speed of the multiple rotors.
  • the number and arrangement of rotors are not limited to those described above or shown in the drawings.
  • the number and arrangement of rotors in Embodiments 7-10 may also be referred to.

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Abstract

一种结合多轴旋翼的自旋翼飞行器，包括机身（4），推进装置其包括推进桨和推进驱动装置（31）；至少一组多轴旋翼系统；每组多轴旋翼系统包括两对称设置于机身两侧的旋翼模块（1）；每个旋翼模块包括至少一个旋翼（11），每个旋翼（11）对应连接有旋翼驱动装置（13）。该飞行器结合了多轴旋翼和自旋翼飞行器，可实现垂直起降、悬停、巡航、左右侧飞等飞行姿态；在作业时可更灵活的选择适宜的飞行姿态，可大幅提高自旋翼飞行器的适用范围和安全性，并且根据多轴旋翼系统安装与否构成不同的构型的飞行器，应用范围大幅提高。

Description

一种结合多轴旋翼的自旋翼飞行器 技术领域
本发明属于飞行器技术领域，具体涉及一种结合多轴旋翼的自旋翼飞行器。
背景技术
自旋翼飞行器是一种以自转旋翼作为升力面、推进桨推/拉力或其它供能方式为前进动力的旋翼类飞行器。自旋翼机的旋翼依靠飞行时前方来流驱动实现自转来提供升力。一旦它发动机空中停车，可以依靠自转旋翼安全着陆，因此历史上它曾被称为“最安全的飞行器”。但由于自旋翼飞行器起飞时需要一定的滑行跑道，无法垂直起降，在低空飞行时如遇到树木、电线杆等障碍物时不能很快拔升，且自旋翼飞行器只能保持一种前行飞行姿态，也无法做到空中悬停和左右侧飞等，从而导致该类自旋翼飞行器适用范围大幅减少，且使用过程中无法及时拔升或快速更改移动方向导致存在诸多安全隐患。而常见的多轴旋翼飞行器，可以解决在狭小场地垂直起降的问题，但由于其固定支架结构的设计，垂直升降时极易受气流影响，且阻力较大，导致垂直升降时不稳定且能耗较大。
为此亟需研发一种兼有多轴旋翼飞行器和自旋翼飞行器优点的飞行器。
但是，通过多轴旋翼和自旋翼简单结合的飞行器还存在诸多缺点，例如该类飞行器的电池舱设置于机身内，不利于大功率电池组的散热和更换等问题，且由于使用范围较窄，不能灵活应对各类使用场景等问题；进一步优化多轴旋翼自旋翼飞行器使其具有更高的稳定性、使用范围更广泛、续航能力更强等也是本领域技术人员亟需解决的问题。
发明内容
本发明为了解决背景技术中所提出的技术问题，提供了一种结合多轴旋翼的自旋翼飞行器。
本发明的技术方案为：
一种结合多轴旋翼的自旋翼飞行器，包括
机身；
自旋翼，设置于所述机身的顶部；
推进装置，设置于机身的尾部，其包括推进桨和驱动所述推进桨转动的推进驱动装置；
至少一组多轴旋翼系统，用于为所述机身提供垂直上升的升力；每组所述多轴旋翼系统包括两对称设置于所述机身两侧的旋翼模块；每个所述旋翼模块包括至少一个旋翼，每个所述旋翼对应连接有旋翼驱动装置，所述旋翼模块通过支架安装于所述机身上；
起落装置，设置于所述机身底部或所述支架上。
进一步优选的，还包括：飞行动力系统，其包括：动力控制模块、燃油供应装置、电力供应装置；
所述动力控制模块分别与所述燃油供应装置、所述推进驱动装置、所述旋翼驱动装置以及所述电力供应装置连接；
根据所选的飞行状态，在垂直起降时控制所述电力供应装置给所述旋翼驱动装置提供电力，驱动所述旋翼转动；在平行飞行时控制燃油供应装置给所述推进驱动装置提供动力，或者控制所述电力供应装置给所述推进驱动装置提供电力，驱动所述推进桨转动。
进一步优选地，所述多轴旋翼系统可拆卸式安装于所述机身；其中，当安装所述多轴旋翼系统时，构成多轴旋翼自旋翼结合的飞行器；当拆卸所述多轴旋翼系统时，构成自旋翼飞行器。
进一步优选地，所述支架包括旋翼架、支撑架，所述旋翼模块通过所述旋翼架安装于所述支撑架上；所述电力供应装置内置于所述支撑架上，或挂载于所述支撑架，或挂载于所述机身外部。
进一步优选的，所述多轴旋翼系统包括至少两个旋翼模块，所述旋翼模块关于所述机身的平均重心对称设置。
进一步优选的，每个所述旋翼模块包括两个旋翼，且两所述旋翼上下同轴设置。
进一步优选的，所述旋翼模块呈八字形对称分布于所述机身两侧。
进一步优选的，每个所述旋翼配备的所述旋翼驱动装置为驱动电机，所述驱动电机设置于所述支架上。
进一步优选的，在所述推进驱动装置为燃油发动机时，所述机身内设有发电机，所述发电机与所述燃油发动机连接，用于给所述电力供应装置充电。进一步优选的，所述支架采用可折叠和/或可伸缩的支架。
进一步优选的，多个所述电力供应装置的重心与飞行器的平均重心位于同一竖向高度。
进一步优选的，所述电力供应装置包括电池模块和电池舱，所述电池舱设置于所述支架下方，用于储放所述电池模块，且所述电池舱通过拆卸组件与所述支撑架可拆卸连接。
进一步优选的，所述支撑架上与所述电池舱对应的位置处设置电池舱加强肋，所述电池舱通过拆卸组件吊挂于所述电池舱加强肋上；所述拆卸组件包括：
连接座，其具有一插口，所述连接座固定于所述支撑架上；
连接头，设置于所述电池舱上，所述连接头与所述插口匹配，所述连接头插设于所述插口内，实现所述连接头的径向限位；
径向插件，沿所述连接座径向插设于所述连接座和所述连接头，实现所述连接头的轴向限位；
固定件，用于限制所述径向插件径向移动脱离所述连接座和所述连接头。
进一步优选的，所述拆卸组件还包括加强框架，所述电池舱和所述旋翼纵梁周侧均围设有所述加强框架；所述连接头设置于所述加强框架上。
进一步优选的，所述支架还包括支撑梁，所述支撑梁一端与所述支架连接，一端与所述机身连接，所述支撑梁与所述机身、所述支撑架连接呈三角结构。
进一步优选的，安装所述多轴旋翼系统时，飞行器载人后的总重量小于或等于飞行器的最大许可起飞重量。
本发明提供了一种结合多轴旋翼的自旋翼飞行器，使其与现有技术相比具有以下的有益效果：
(1)本发明中创造性的将电力供应装置吊挂设置于飞行器外侧，其优点在于：一、无需在机身内另外设置容纳空间，避免占用机身内部本就局促的空间，也无需额外加长机身扩大飞行器内部空间，避免增加飞行器的设计难度和生产成本；二、电力供应装置外挂于飞行器外侧有利于快速拆卸更换，飞行器到达目的地后可直接更换备用电池舱内的电池模块或者直接更换电力供应装置，无需等待充电，有效提高工作效率；三、旋翼的旋翼驱动装置工作时电力供应装置高功率放电同时高功率放热，电力供应装置设置于机身内部不利于散热，为电力供应装置布置液冷系统又会增加重量并增加系统的复杂性；本发明通过将电力供应装置外挂，电力供应装置直接暴露于空气中，利用风冷提高电力供应装置的散热能力，当飞行器不航行时，电力供应装置或者电池舱内的电池模块可取出储放于室内，避免环境温度非常低时电量损耗严重，提高电池寿命，从而机身内部无需额外配备冷却系统和保温系统。
(2)在本发明中，多轴旋翼系统可拆卸且优选的飞行器为载人飞行器；当安装多轴旋翼系统时，构成多轴旋翼自旋翼结合的载人飞行器；当拆卸多轴旋翼系统后，构成的自旋翼飞行器仍然可以载人，满足载人飞行器的适航标准及规范，扩大飞行器的应用范围。
(3)在本发明中，可根据不同最大起飞重量的应用场景，更换适配构型的自旋翼，有效扩大飞行器的应用范围。
(4)在本发明中，将多个多轴旋翼系统集成设置于自旋翼飞行器，通过多个多轴旋翼系统与自旋翼、推进装置的配合可实现垂直起降，空中悬停，空中巡航、左右侧飞等多种飞行姿态；从而本方案中的自旋翼飞行器无需跑道也可起飞，在作业时可更灵活的选择适宜的飞行姿态，从而可大幅提高复合式自旋翼飞行器的适用范围以及使用的安全性。
(5)相比现有固定翼垂直起降飞行器，本发明同样能够实现垂直起降、悬停等飞行姿态，然而由于其固定翼的结构，飞行器整体的结构更加复杂，体积更大，垂直升降时极易受气流影响，且阻力较大，导致垂直升降时飞行器不稳定且能，影响飞行姿态调控；且其起飞需要增加辅助装置，从而其安全性、可靠性更低，此外其经济性与气动效率有所降低，耗能更大；而本发明提供的结合多轴旋翼的自旋翼飞行器，不仅能够完美完成垂直起降等多种飞行姿态，同时摒弃固定翼垂直起降飞行器的上述缺点；具体的其结构相比于固定翼垂直起降飞行器更简单，体积更小，从而可大幅降低结构成本以及维修维护成本；更重要的是，由于其多轴旋翼系统的设置，其飞行阻力更小，不易受到气流的影响，飞行更加安全，通过多组多轴旋翼系统不同的排布，增加其飞行可靠性；且起飞时无需借助其他辅助设置，经济性与气动效率更高；
(6)本发明中，多轴旋翼系统具有多种排布方式，例如呈八字形排布或呈环形排布，从多个自由度调节飞行器飞行姿态，有效提高对飞行器飞行姿态的控制，提高飞行器飞行的稳定性和灵活性，减少气流的影响，适应更广的飞行环境。
(7)相比于设置有尾翼的机身，本发明给出了无尾翼设计，省去了尾翼结构，在调控飞行器的姿态时，通过调控多个旋翼的转速实现偏转和俯仰等飞行器飞行姿态的改变，而无需通过尾翼实现飞行姿态的调控；省去尾翼有利于减小机身结构的重量和体积，在飞行器飞行时也有利于减小空气阻力，提高旋翼调控飞行姿态时的反应灵敏度。
附图说明
结合附图，通过下文的述详细说明，可更清楚地理解本发明的上述及其他特征和优点，其中：
图1为本发明实施例1中结合多轴旋翼的自旋翼飞行器的示意图；
图2为本发明实施例1中结合多轴旋翼的自旋翼飞行器的俯视图；
图3为本发明实施例1中电池舱安装于支架上的示意图；
图4为本发明实施例1中电池舱的俯视示意图；
图5为本发明实施例1中拆卸组件的结构示意图；
图6为本发明实施例2中结合多轴旋翼的自旋翼飞行器的示意图；
图7为本发明实施例3中结合多轴旋翼的自旋翼飞行器的示意图；
图8为本发明实施例4中结合多轴旋翼的自旋翼飞行器的示意图；
图9为本发明实施例4中结合多轴旋翼的一种自旋翼飞行器的示意图；
图10为本发明实施例4中结合多轴旋翼的自旋翼飞行器的俯视图；
图11为本发明实施例5中结合多轴旋翼的自旋翼飞行器的示意图；
图12为本发明实施例6中结合多轴旋翼的自旋翼飞行器的示意图；
图13为本发明实施例7中结合多轴旋翼的自旋翼飞行器的示意图；
图14为本发明实施例8中结合多轴旋翼的自旋翼飞行器的示意图；
图15为本发明实施例9中结合多轴旋翼的自旋翼飞行器的示意图；
图16为本发明实施例10中结合多轴旋翼的自旋翼飞行器的示意图；
图17为本发明实施例11中结合多轴旋翼的自旋翼飞行器的示意图；
图18为本发明实施例1中结合多轴旋翼的自旋翼飞行器垂直飞行原理图；
图19为本发明实施例1中结合多轴旋翼的自旋翼飞行器不同姿态的飞行原理示意图。
符号说明：
1-旋翼模块；11-旋翼；12-旋翼架；13-旋翼驱动装置；14-防护罩；2-拆卸组件；21-连接座；22-加强框架；23-连接头；24-径向插件；25-固定件；31-推进驱动装置；3-推进桨；4-机身；5-自旋翼；6-起落装置；7-电力供应装置；8-电池舱；9-支撑架；10-支撑梁；7a-环形支架；7b-弧形支架；15-尾翼。
具体实施方式
为了更清楚地说明本发明实施例或现有技术中的技术方案，下面将对照附图说明本发明的具体实施方式。显而易见地，下面描述中的附图仅仅是本发明的一些实施例，对于本领域普通技术人员来讲，在不付出创造性劳动的前提下，还可以根据这些附图获得其他的附图，并获得其他的实施方式。
为使图面简洁，各图中只示意性地表示出了与本发明相关的部分，它们并不代表其作为产品的实际结构。另外，以使图面简洁便于理解，在有些图中具有相同结构或功能的部件，仅示意性地绘示了其中的一个，或仅标出了其中的一个。在本文中，“一个”不仅表示“仅此一个”，也可以表示“多于一个”的情形。
实施例1
参照图1-5，本发明提供了一种结合多轴旋翼的自旋翼飞行器，包括：
机身4；自旋翼5，设置于机身4的顶部；推进装置，设置于机身4的尾部，其包括推进桨3和驱动推进桨3转动的推进驱动装置31；至少一组多轴旋翼系统，用于为机身4提供垂直上升的升力；每组多轴旋翼系统包括两对称设置于机身4两侧的旋翼模块1；每个旋翼模块1包括至少一个旋翼11，每个旋翼11对应连接有旋翼驱动装置13，旋翼模块1通过支架安装于机身4上；起落装置6，设置于机身4底部或支架上。在本实施例中的起落装置6可以采用前三点式不可收放起落架，前轮可由脚蹬操纵实现地面转弯，主轮带有同步刹车。
优选地，本实施例中的自旋翼设置在机身顶部，并向后倾斜约4°，使自旋翼的旋翼叶片始终与迎风角成约4°。进一步说明，如果自旋翼逆时针旋转，由于倾斜4°，使得机身右侧的叶片总是前缘向前，由于叶片具有一定翼型，产生的阻力较小；机身左侧的叶片总是后缘向前，叶片整体与来流产生较大迎角，可以产生的升力较大。推进桨设置在机身尾部，为飞行器的平飞前进提供动力。
进一步说明，本实施例中的多轴旋翼系统包括对称设置在机身两侧的旋翼模块1，每个旋翼模块1至少一个旋翼11。本实施例中通过改变各个旋翼11的转速来改变飞行器的飞行姿态。
本实施例中的多轴旋翼系统可拆卸式安装于机身4；其中，当安装多轴旋翼系统时，构成多轴旋翼自旋翼结合的飞行器；当拆卸多轴旋翼系统时，构成自旋翼飞行器。进一步说明，本实施例通过可重构的设计理念，根据飞行器的设计需求，改预定好的飞行器构件改装为所需功能的飞行器，即，本实施例中经过重新设计的可重构轻型飞行器，通过不同结构的多轴旋翼系统组成垂直起降体系，成为既能够垂直起降又能跑道起降的混合轻型飞行器，并且它通过拆卸掉多轴旋翼系统就恢复成为自旋翼飞行器。进一步优选说明，本实施例中的可重构飞行器应用于载人飞行器上，使得在安装多轴旋翼系统时，构成多轴旋翼自旋翼结合的载人飞行器。当拆卸多轴旋翼系统时，构成自旋翼载人飞行器。
另外，本实施例中的机身尾部可以设置尾翼15，也可以直接采用无尾翼机身。当机身尾部设置有尾翼15时，尾翼15可以采用平尾或V型尾翼或其他构型的尾翼，此处不做限制。当机身为无尾翼机身时，无尾翼机身上设置有多组多轴旋翼系统。
可以看出，机身4的顶部设置有自旋翼5，底部设置有起落装置6，本实施例中的推进桨3不局限于设置在机身4尾端，在其他实施例中，也可以设置于机身4头部。两个关于机身4纵向对称设置多轴旋翼系统。本实施例中的多轴旋翼系统可拆卸的安装于机身4上，便于改装为所需的飞行器构型。
基于上述本实施例中结合多轴旋翼的自旋翼飞行器的结构，还包括：飞行动力系统，其至少包括：动力控制模块、燃油供应装置、电力供应装置7。动力控制模块分别与燃油供应装置、推进驱动装置31、旋翼驱动装置13以及电力供应装置7连接；根据所选的飞行状态，在垂直起降时控制电力供应装置7给旋翼驱动装置13提供电力驱动旋翼11转动；在平行飞行时控制燃油供应装置给推进驱动装置31提供动力驱动推进桨3转动，或者控制电力供应装置7给推进驱动装置31提供电力驱动推进桨3转动。
本实施例中的飞行动力系统包括至少一组电力供应装置7，并且关于整个飞行器的平均重心对称布置。本实施例中的电力供应装置7可内置于所述机身4内侧，或者挂载于支架或挂载于机身4外部。当然优选地，电力供应装置7挂载于支架或机身4外部。进一步说明，本实施例中的动力控制模块、燃油供应装置、推进驱动装置31设于机身4内部。在一种实施例中，支架包括旋翼架12、支撑架9，旋翼模块1通过旋翼架12安装于支撑架9上。此时优选地，电力供应装置7挂载于支撑架9或机身4外部。当然，进一步优选地，电力供应装置7吊挂于支撑架下方。
本实施例中创造性的将电力供应装置7吊挂设置于飞行器外部，其优点在于：一、无需在机身内另外设置电力供应装置占用机身内部本就局促的空间，也无需额外加长机身扩大飞行器内部空间，避免增加飞行器的设计难度和生产成本；二、电力供应装置外挂于飞行器外部有利于快速拆卸更换，飞行器到达目的地后可直接更换备用电力供应装置，无需等待充电，有效提高工作效率；三、旋翼的旋翼驱动装置工作时利用电力供应装置高功率放电同时高功率放热，而当电力供应装置设置于机身内部不利于电池散热，为电力供应装置布置液冷系统又会增加重量并增加系统的复杂性；本发明通过将电力供应装置外挂，电力供应装置直接暴露于空气中，利用机外冷空气提高电力供应装置的散热能力，当飞行器不航行时，电力供应装置可取出储放于室内，避免环境温度非常低时电力供应装置内电量损耗严重，提高使用寿命，从而机身内部无需额外配备冷却系统和保温系统。
参阅图1，本实施例中的多轴旋翼系统包括至少两个旋翼模块1，即机身4每一侧各有两个旋翼模块1，使得在调控飞行器的飞行姿态时可从四个自由度进行调控，增强飞行器飞行的稳定性和灵活性。当然在其他实施例中，旋翼模块1设置的组数不局限于以上所述，可根据飞行器的构型以及所需的垂直起降的升力大小、调控的稳定性灵敏性等调整旋翼模块1的个数。例如多轴旋翼系统包括四个旋翼模块1或五个旋翼模块1等。进一步优选的，位于机身4同一侧的旋翼模块1两两共用一支架。优选地，位于机身4同一侧的旋翼模块1两两共用一旋翼架12，两旋翼模块1分别设置于旋翼架12的两端且位于同一水平面上，特别注意的是，旋翼模块1关于机身4的平均重心对称设置，且两多轴旋翼系统中对应的两旋翼模块1关于机身纵向相互对称，从而使得机身4两侧和机身4前后方向的旋翼模块1两两相互对称，拉力基本相当，飞行器受力平衡，从而无需损耗过多的能量用于调节各旋翼11之间的平衡，大大减少了能量的损耗。当然在其他实施例中，每个多轴旋翼系统内的两旋翼模块1的相对位置不局限于以上所述，可根据机身的结构和实际应用做相应的布局调整。当然在其他实施例中，也可以是每个旋翼模块对应一旋翼架。
进一步说明，当多轴旋翼系统包括两个旋翼模块1，每个旋翼模块1上包括一组旋翼，使得整个飞行器包括四个旋翼模块，进一步包括四组旋翼，每个旋翼连接一旋翼驱动装置，参见图18所示，假设旋翼驱动装置分别为A、B、C、D，当飞行器进行垂直运动时，相邻的两组的旋翼转动方向相反，而在对角线上的两个旋翼驱动装置的驱动旋翼转动方向相同，且转速相等。进一步说明，为了保证旋翼转动产生的升力都是向上的，一组对角线上的旋翼驱动装置需要驱使旋翼顺时针方向转动以便能产生向上升力；另一组对角线上的旋翼驱动装置需要驱使旋翼进行逆时针方向的旋转产生向上升力。也可以理解为，各组旋翼在产生升力时，气流往下吹出，以提供向上拉力。
进一步说明含四个旋翼模块1的飞行器的飞行姿态控制原理。
参见图19(a)，飞行器垂直运动时，同时增加四个旋翼驱动装置的输出功率﹐旋翼转速增加使得总的拉力增大，当总拉力足以克服整机的重量时，飞行器便离地垂直上升；反之，同时减小旋翼驱动装置的输出功率，飞行器则垂直下降，直至平衡落地，实现了沿z轴的垂直运动；当外界扰动量为零时，在旋翼产生的升力等于飞行器的自重时，飞行器便保持悬停状态。
参见图19(b)，在飞行器俯仰运动时，旋翼驱动装置A的转速上升，旋翼驱动装置C的转速下降，转速的改变量大小应相等，旋翼驱动装置B、旋翼驱动装置D的转速保持不变。由于旋翼驱动装置A驱动的旋翼11产生的升力上升，旋翼驱动装置C驱动的旋翼11升力下降，产生的不平衡力矩使机身绕y轴旋转，同理，当旋翼驱动装置A的转速下降，旋翼驱动装置C的转速上升，机身便绕y轴向另一个方向旋转，实现飞行器的俯仰运动。
参见图19(c)，当飞行器滚转运动时，改变旋翼驱动装置B和旋翼驱动装置D的转速，保持旋翼驱动装置A和旋翼驱动装置C的转速不变，则可使机身绕x轴旋转(正向或反向)，实现飞行器的滚转运动。
参见图19(d)，当飞行器偏航运动时，当旋翼驱动装置A和旋翼驱动装置C的转速上升，旋翼驱动装置B和旋翼驱动装置D的转速下降时，旋翼驱动装置A和旋翼驱动装置C对机身的反扭矩大于旋翼驱动装置B和旋翼驱动装置D对机身的反扭矩，机身便在富余反扭矩的作用下绕z轴转动，实现飞行器的偏航运动，转向与旋翼驱动装置A、旋翼驱动装置D的转向相反。
参见图19(e)，当飞行器前后运动时，增加旋翼驱动装置C转速，使拉力增大，相应减小旋翼驱动装置A转速，使拉力减小，同时保持其它两个电机转速不变，反扭矩仍然要保持平衡。飞行器首先发生一定程度的倾斜，从而使旋翼拉力产生水平分量，因此可以实现飞行器的前飞运动。向后飞行与向前飞行正好相反。
参见图19(f)，当飞行器侧向飞行时，基于飞行器结构对称原则，所以飞行器侧向飞行的工作原理与前后运动的工作原理一致。
进一步阐述，当飞行器中的多轴旋翼系统采用其他实施例中的构型时，各旋翼控制可以采用上述的飞行姿态控制原理。对此本实施例不再进一步阐述。
进一步的，在本实施例中，参阅图1，优选的旋翼模块1包括两个旋翼11，且两旋翼11上下同轴设置，形成共轴式双旋翼的方式，大幅提升每个推进桨模块的总功率，其垂直起降升力变强的同时结构更紧凑，其占用的空间体积相更小，重量效率更高；进一步的，共轴双旋翼的方式使得在悬停状态下两同轴旋翼间的气动干扰会产生有利影响，提高悬停效率。
在本实施例中，参阅图1-3，优选地，电力供应装置7挂载于支撑架9下方。此时电力供应装置7的重力方向和支撑架9受到的升力方向相反，能够抵消支撑架9和机身4 上的的一部分气动弯矩，因此，支撑架9挂载重物后，其强度反而过剩，因此对支撑架9的强度要求可以适当降低，且对机身也无需进行额外的补强设计，简化设计难度。当然在其他实施例中，电力供应装置7挂载的位置不局限于支撑架9，例如也可以挂载于机身外部。
在本实施例中，支撑架9内部的承力梁与机身内部的梁螺栓连接，在机舱内即可完成支撑架9与机身4的连接，拆卸方便且利于检修。当然也可以采用插销连接结合螺栓连接的方式；当然可以通过在机身上设置接头，支撑架通过接头与机身连接，此处不做限制。
进一步优选的，在本实施例中，参阅图2，支撑架9一端与机身4连接，另一端沿机身纵向向机身头部前倾，形成前掠式布局，如此支撑架9可避开飞行器机舱门的位置布置，且使得位于机身同一侧的两旋翼模块能够关于机身的平均重心对称，并且使支撑架9下方挂载的电力供应装置7尽可能的靠近机身的平均重心纵向，整机重量分布更均匀，大大提高了飞行器的稳定性和平衡。当然在其他实施例中，支撑架9设置的方向以及位置不局限于以上所述或图中所示，需根据机身的实际布局进行对应的调整，避免支撑架9阻碍进出飞行器且保障旋翼模块和电力供应装置7布局合理。例如支撑架9也可以是沿机身对称设置的圆圈形支架或两个相向倾斜呈八字型的支架。
进一步优选的，参阅图2，本实施例中的支撑架9包括两个电力供应装置7，两个电力供应装置7沿支撑架9的长度方向排布。在支撑架9为前掠式布局的基础上，两电力供应装置7沿支撑架9的延伸方向设置，使得电力供应装置7尽可能的靠近机身的平均重心纵向，整机重量分布更均匀，大大提高了飞行器的稳定性和平衡。当然在其他实施例中，电力供应装置7的布局不局限于以上所述。进一步优选的，多个电力供应装置7的重心与飞行器的平均重心位于同一竖向高度，使得在挂载电力供应装置7时飞行器的平均重心不存在竖直方向上的移动。另外，当支架中的支撑架9与旋翼架12位于同一个时，电力供应装置7沿支撑架9/旋翼架12的长度方向排布，其他均与支撑架原理相同，此处不再赘述。
参阅图1-5，当电力供应装置7挂载于支撑架9或机身4外部时，本实施例中的电力供应装置7包括电池模块和电池舱8，电池舱8设置于支撑架9下方，用于储放电池模块，且电池舱8通过拆卸组件与支撑架9可拆卸连接。通过设置电池舱8，便于快速的安装电池模块，且电池舱8内还可以设置冷却或保温装置用于对加快电力供应装置7的冷却或保温速度。本实施例中的电池模块包括但不局限于采用锂电池、蓄电池。
进一步优选的，支撑架9上与电池舱8对应的位置处设置电池舱加强肋，电池舱9通过拆卸组件2吊挂于电池舱加强肋上，避免支撑架9局部受力过大，从而延长支撑架9的使用寿命。
参阅图3-5，本实施例中的拆卸组件2包括：连接座21，其具有一插口，连接座21固定于支撑架9上；连接头23，设置于电池舱8上，连接头23与插口匹配，连接头23插设于插口内，实现连接头的径向限位；径向插件24，沿连接座21径向插设于连接座21和连接头23，实现连接头的轴向限位；固定件25，用于限制径向插件24径向移动脱离连接座21和连接头 23。通过拆卸结构可实现电池舱8以及电力供应装置7的快速拆卸和更换。当然在其他实施例中，拆卸组件的具体结构不局限于以上所述或图中所示，例如也可以是其他连接组件。
进一步优选的，在本实施例中，参阅图4、5，拆卸组件2还包括加强框架22，电池舱8周侧围设有加强框架22；连接头23设置于加强框架22上，加强框架用于提高电池舱8的强度。
在本实施例中的径向插件24和固定件25均为插销结构，当然在其他实施例中，径向插件和固定件的具体结构不局限于以上所述，例如也可以是其他连接件或螺母、螺栓结构等。
在本实施例中，参阅图3，优选的支撑架9的横截面为椭圆流线型截面，其底部可提供更大的面积用于悬挂电力供应装置7，椭圆流线型截面的流线型为前圆后尖，略像水滴形状，从而支撑架9对电力供应装置7的支撑更稳定，强度更高；且流线型的支撑架9既能产生升力补充飞行器的升力，又不会增大空气阻力。当然在其他实施例中，支撑架9的截面形状不局限于以上所述或图中所示，例如支撑架9的横截面也可以是圆形。
在本实施例中，结合多轴旋翼的自旋翼飞行器中的飞行动力系统包括动力控制模块、电力供应装置7、燃油动力装置，动力控制模块可控制电力供应装置7和燃油动力装置一起输出动力。其中电力供应装置7和燃油动力装置均可为飞行器提供动力，两者的动力可共同输出或单独输出至对应的装置。例如电力供应装置7为旋翼驱动装置13提供能量，从而旋翼驱动装置13驱动旋翼转动，实现飞行器垂直起降、多轴模式的低速飞行等，而燃油动力装置的燃油发动机为推进驱动装置31提供动力，驱动推进装置中的推进桨3转动，为飞机的滑跑起降和正常飞行提供前进的动力。当然电力供应装置7和燃油动力装置也可以共同为旋翼驱动装置13、推进驱动装置31提供动力。当然在其他实施例中，结合多轴旋翼的自旋翼飞行器的动力全部来源于电力动力系统，控制系统控制电力动力装置为旋翼驱动装置13、推进驱动装置31提供动力，其中推进驱动装置31的发动机为电动发动机。电力动力系统的动力来源于以上所述的电力供应装置7和/或额外设置的电池组，无需设置燃油动力装置。
进一步说明，本实施例中的每个旋翼11配备的旋翼驱动装置13为驱动电机，驱动电机设置于支架(可以在支撑架或者旋翼架)上。当然驱动电机也可以设置在其他位置，此处不做限制，可根据实际情况选择最佳的设置位置。
本实施例中的推进驱动装置31可以为燃油发动机或电力发动机，当推进驱动装置为燃油发动机时，通过燃油供应装置提供动力，当推进驱动装置31为电力发动机时，通过电力供应装置7提供电力。进一步地，在推进驱动装置为燃油发动机时，机身内设有发电机，发电机与燃油发动机连接，用于给电力供应装置7充电，在飞行过程中，可以给电力供应装置7充电，确保垂直升降时能源的供应。值得说明的是，当推进驱动装置31为燃油发动机时，发动机启动需要预热，推进驱动装置31预热时推进桨3转动产生一定的推力使得机向前滑行，此时需用制动装置刹住机身使得机身在地面时不向前滑行；当飞行器开始远离地面起飞时，此时机身4前后两组多轴旋翼系统的转速不一致，进而施加于机身 4一定的向后的力与推进桨3产生的推力相抵消，从而确保飞行器能够垂直起飞。从而当推进驱动装置31为燃油发动机时，需设置至少两组多轴旋翼系统。当推进驱动装置31为电发动机时，无需预热，推进桨3可在垂直起飞后再启动，从而不会对机身4产生向前的推力，此时只需设置至少一组多轴旋翼系统。
在本实施例中优选的支架(可以为支撑架、旋翼架或者支撑架和旋翼架)为可折叠支架，当飞行器未作业时，支架折叠放置，节省其存放空间。
在本实施例中，优选的支架(可以为支撑架、旋翼架或者支撑架和旋翼架)为可伸缩支架，从而可调节支架的长度以获得最佳的飞行效果；或当飞行器未作业时，可收缩支架，节省其存放空间。当然在其他实施例中，直接支架也可以是可折叠且可伸缩的支架，此处不做限制。
在本实施例中，通过调控多组旋翼的转速、推进桨的转速，使得结合多轴旋翼的自旋翼飞行器实现多种飞行姿态，以下举例说明飞行器实现多种飞行姿态的过程。垂直起飞状态时，起飞前发动机预热后，同时启动多个旋翼11的旋翼驱动装置13，多个旋翼11同时转动并产生足够的升力，从而飞行器实现垂直起飞；需注意的是，推进驱动装置31预热时推进桨3转动产生一定的推力使得机身向前滑行，当飞行器开始起飞并与地面无接触时，制动装置无法提供阻力阻止机身前行，此时调节机身前后两组旋翼11的转速，使其转速不一致，进而施加于机身一定的向后的力与推进桨3产生的推力相抵消，从而确保飞行器能够垂直起飞。巡航状态时，随着推进驱动装置31的持续工作，推进桨3向前的推力增大，推动飞行器前飞，自旋翼5通过前飞来流驱动自转，产生升力，随着飞行器前飞速度进一步加快，这一升力迅速提高，多个旋翼11适时降低转速，提供的升力随之逐步降低，当前飞时自旋翼5桨叶产生的升力足够支撑飞行器水平飞行时，多个旋翼11完全停止，飞行器完全靠自旋翼5产生的升力保持稳定高度的水平飞行。悬停状态时，升高旋翼11的转速同时降低推进驱动装置31的转速，直至旋翼11产生的升力等于飞行器自重时，保持旋翼11转速不变，从而飞行器保持悬停状态。在悬停状态下，还可通过调整机身两侧的旋翼11的转速，实现飞行器左右侧飞。当需要降落时，逐步降低旋翼11转速使飞行器平稳垂直降落，也可以像自旋翼飞机那样在跑道上滑行降落。当然，结合多轴旋翼的自旋翼飞行器还可实现倒飞运动、俯仰运动等，此处不一一详述。
在本实施例中，优选的飞行器为载人飞行器，参阅图1，进一步说明，当安装有多轴旋翼系统时时，飞行器构成多轴旋翼自旋翼结合的载人飞行器；当然在其他实施例中，拆卸多轴旋翼系统后，飞行器构成自旋翼飞行器，值得注意的是，该自旋翼飞行器仍然可载人。可根据使用需求选择合适的构型，扩大飞行器的应用范围。其中无论多轴旋翼系统是否安装，均需确保构成的飞行器均可载人，均满足载人飞行器的适航审定标准，从而相比于现有技术，本发明通过对机身进行减重设计，降低飞行器的空机重量，使得本发明中的自旋翼飞行器的空机重量远小于现有技术中的自旋翼飞行器的空机重量，以确保飞行器强度满足两种构型转换下仍能满足载人要求。
进一步的，在本发明中，安装多轴旋翼系统时，飞行器载人后的总重量仍然小于或等于飞行器的最大许可起飞重量，确保该飞机在任何构型下满足载人飞机的适航标准。在本 实施例中，可根据飞行器的最大许可起飞重量更换不同构型的自旋翼。例如针对于最大起飞重量较轻的飞机器需换装较小的自旋翼。根据应用场合适配对应自旋翼，灵活变化机型，使用范围更广。其中需注意的是，不同构型的自旋翼与机身的接口相匹配，确保不同构型的自旋翼可换装于同一机身上。
实施例2
本实施例是基于实施例1的基础上做的改进，参阅图6，本实施例相比于实施例1的区别在于：
在本实施例中，支架还包括支撑梁10，支撑梁10一端与支撑架9连接，一端与机身4连接，支撑梁10与机身4、支撑架9连接呈三角结构，有效的增强了支架的强度。具体的，在本实施例中，参阅图6，支撑梁一端与机身内部的自旋翼的承力梁连接，一端与支撑架9连接呈三角结构，当然在其他实施例中，支撑梁也可以与机身的主传力梁连接呈三角结构，即支撑梁10和支撑架9位于同一水平面。
本实施例中的其他结构及其连接关系均与实施例1相同，此处不再赘述。
实施例3
本实施例是基于实施例2的基础上做的改进，参阅图7，本实施例相比于实施例1的区别在于：在本实施例中，支撑架9的横截面为圆形。在飞行器前进时起飞时，横截面为椭圆流线型的支架，也会产生升力，对整机的平衡和稳定造成一定的影响，从而本实施例中通过设置圆柱状的支撑架9，支撑架9受风力的面积变小，从而飞行时支撑架9受到的升力变小甚至消失，减少对飞行器平衡性和稳定性的影响，有利于更准确的操控飞行器飞行。本实施例中的其他结构及其连接关系均与实施例1相同，此处不再赘述。
实施例4
本实施例是在实施例1的基础上做的改进，本实施例与实施例1的区别在于：
在本实施例中，参阅图8-10，多轴旋翼系统包括四个旋翼模块1，两两旋翼模块1共用一旋翼架12和一支撑架9，其中两旋翼模块1设置于机身头部，另外两旋翼模块1设置于机身4平均重心位置处，设置于机身4头部的两旋翼模块1中的旋翼的长宽小于另外两旋翼模块1中的旋翼11，从而在进一步提高飞行器升力的同时可有效降低机身头部的旋翼模块对机身的平均重心的影响。进一步地，在本实施例的多轴旋翼系统其中位于机身同一侧的四个旋翼模块1，两两共用一个支撑架9，且四组旋翼模块1分别设置于机身4的头部和尾部两端，从四个自由度调控飞行器的飞行姿态，增强飞行器飞行的稳定性和灵活性。旋翼模块1包括一个旋翼11，当然在其他实施例中，旋翼模块1包括的旋翼11的数量不局限于以上或图中所示，例如旋翼模块1也可以包括两个旋翼11。每个旋翼11配备的旋翼驱动装置13为驱动电机，驱动电机内置于支架(可以是支撑架、旋翼架)上。当然驱动电机也可以设置在其他位置，此处不做限制，可根据实际情况选择最佳的设置位置。
本实施例提供的结合多轴旋翼的自旋翼飞行器，通过多轴旋翼系统与自旋翼、推进装置的配合可实现垂直起降，空中悬停，空中巡航、左右侧飞等多种飞行姿态；从而本实施例提供的结合多轴旋翼的自旋翼飞行器无需跑道也可起飞，在作业时可更灵活的选择适宜的飞行姿态，从而可大幅提高自旋翼飞行器的适用范围以及使用的安全性。
本实施例中的其他结构及其连接关系均与实施例1相同，此处不再赘述。
实施例5
本实施例是在实施例1的基础上做的改进，本实施例与实施例1的区别在于：
在本实施例中，参阅图11，旋翼模块1还包括防护罩14，防护罩14围设于旋翼周侧，从而在不影响旋翼竖向气流流动的同时，可对旋翼11起到保护作用，防止旋翼11转动时误伤的人或物，当然也可防止高速转动时旋翼11松动脱落而打到机身上等。在本实施例中，每个旋翼11配备一防护罩14，防护罩14的横截面为圆形围设于旋翼11周侧，其中防护罩14设置于旋翼架12上。当然在其他实施例中，防护罩14的具体形状不局限于以上所述或图中所示。例如防护罩14横截面也可以是椭圆形的防护罩。
本实施例中的其他结构及其连接关系均与实施例1相同，此处不再赘述。实施例6
本实施例是基于实施例1的基础上做的改进，参阅图12，本实施例相比于实施例1的区别在于：
在本实施例中，支撑架9为机翼结构，在飞行器滑行起降时，机翼结构进一步的为飞行器提供升力，使得对旋翼的功率要求更低。当然在其他实施例中，支架的具体结构不局限于以上所述或图中所示。
进一步的，在本实施例中，多轴旋翼系统包括四个旋翼模块1，两旋翼模块1共用一旋翼架12，且该两旋翼模块1关于机身的平均重心对称设置，大大提高了飞行器的稳定性和平衡。
本实施例中的其他结构及其连接关系均与实施例1相同，此处不再赘述。
实施例7
参阅图13，本实施例是基于实施例1、4的基础上做的改进，每个旋翼模块1包括两个旋翼11，且两旋翼11同轴上下设置，形成共轴式双旋翼的方式，共轴双旋翼的方式使得每个旋翼模块1的总功率更大，且体积小，结构紧凑、重量效率高。具体地，本实施例中设置两组多轴旋翼系统，每个旋翼模块1通过一支架(包括支撑架9和旋翼架12)安装于机身4上，旋翼模块1包括两个旋翼11，且两旋翼11同轴上下设置同一支架上。当然在其他实施例中，对多轴旋翼系统设置的组数不做限制，例如也可以设置三组多轴旋翼系统。
本实施例中的其他结构及其连接关系参见实施例1、4，此处不再赘述。
相比于实施例4中的每个旋翼模块1包括一个旋翼11且设置四组旋翼模块1，位于机身4同侧的旋翼11两两共用一个支撑架9的方式，本实施例中旋翼模块1使用共轴双旋翼的方式，大幅提升每个旋翼模块1的总功率，其垂直起降升力变强的同时结构更紧凑，其占用的空间体积相比实施例4几乎减小了一倍，重量效率更高；进一步的，共轴双旋翼的方式使得在悬停状态下两同轴旋翼间气动干扰会产生有利影响，提高悬停效率。
实施例8
参阅图14，本实施例是基于实施例1、4的基础上做的改进，本实施例与实施例1、4的区别在于：
在本实施例中，设置两组以上的多轴旋翼系统，旋翼模块1绕机身4周向均布设置。从而旋翼11对飞行器的升力更加均衡，从而飞行器飞行或起降时更加平稳。本实施例中的其他结构及其连接关系参见实施例1，此处不再赘述。
实施例9
本实施例是在实施例4的基础上做的更改，本实施例与实施例4的区别在于：
在本实施例中，参阅图15，旋翼模块1呈八字形对称分布于机身4两侧。具体的，在本实施例中，设置四组多轴旋翼系统，位于机身两侧的四个旋翼模块1沿机身轴向排布呈八字形。在飞行器飞行前进过程中，多轴旋翼系统呈八字形排布，从而增大机身尾部旋翼11与推进桨3的距离，从而两者同时转动时，两者周围产生的气流无相互扰动或干扰较小，从而有利于提高气动效率。
进一步地，呈八字形排布的多组旋翼通过两支撑架连接，两支架下侧面对称设置有四个起落装置。优选地，位于机身同一侧的旋翼模块共用一支撑架，两侧的支撑架呈八字形排布。本实施例中的两支架下侧面对设置有四个起落装置。四个起落装置设置于八字形排布的两支架下方，形成类矩形排布，在飞行器降落时提供更平衡稳定的支撑力，从而飞行器降落更平稳，提高用户舒适度。
实施例10
本实施例是在实施例4的基础上做的更改，本实施例与实施例4的区别在于：
参阅图16，在本实施例中，设置四组多轴旋翼系统，其中两组多轴旋翼系统环绕机身1周向设置，且四个旋翼模块通过一环形支架连接形成环形部；另外两组多轴旋翼系统设置于环形部外侧，且位于同一侧的两旋翼模块通过一弧形支架连接于环形部两侧。进一步的，多轴旋翼系统设置与机身顶部，当然在其他实施例中，对多轴旋翼系统设置的位置不做限制。例如多轴旋翼系统也可以设置于机身中部或机身底部。
本实施例相比于实施例4，其优点在于，本实施例中螺旋组件的排布有利于提高飞行器的稳定形，便于调控各个旋翼的转速从而调控飞行器飞行的状态。
实施例11
本实施例是在实施例1的基础上做的更改，本实施例与实施例1的区别在于：
在本实施例中，参阅图8、9、17，机身为无尾翼机身，且无尾翼机身上至少设置有两组多轴旋翼系统。当机身设置有尾翼11时，在巡航前飞时通过尾翼控制飞行器完成偏转或俯仰等运动时，此时旋翼处于关闭状态。当省去尾翼时，飞行器飞行时需持续开启至少两组多轴旋翼系统，通过调控多个旋翼的转速差，调控飞行器的飞行姿态。
以下举例说明无尾翼情况下飞行器改变姿态的调控过程：
偏转运动时，旋翼转动过程中由于空气阻力作用会形成与转动方向相反的反扭矩，为了克服反扭矩影响，可使位于对角线上的各个旋翼转动方向相同，具体的，参阅图16，可使旋翼A2、C2、B1、D1顺时针旋转，对应的A1、C1、B2、D2逆时针旋转；反扭矩的大小与旋翼转速有关，当四组旋翼转速相同时，产生的反扭矩相互平衡，飞行器不发生转动，当多个旋翼的转速不完全相同时，不平衡的反扭矩会引起飞行器偏转；当旋翼A2、C2、B1、D1的转速上升，旋翼A1、C1、B2、D2转速下降，机身便在反扭矩的作用下转动，实现飞行器的偏转运动，转向与旋翼A2、C2、B1、D1的转向相反。
俯仰运动时，机身前后的旋翼转速不一致，则产生不平衡的扭矩使得机身发生俯仰运动；具体的，旋翼A1、A2、B1、B2转速上升，C1、C2、D1、D2转速下降，且每个旋翼的转速改变量相等，由于旋翼A1、A2、B1、B2的升力上升，C1、C2、D1、D2升力下降，产生的不平衡力矩使机身前翘；同理，旋翼A1、A2、B1、B2转速下降，C1、C2、D1、D2转速上升，机身便后翘，从而可实现飞行器的俯仰运动。
相比于设置有尾翼的机身，本实施例中无尾翼的机身，省去了尾翼结构，在调控飞行器的姿态时，通过调控多个旋翼的转速实现偏转和俯仰等飞行器飞行姿态的改变，而无需通过尾翼实现飞行姿态的调控；有利于减小机身结构的重量和体积，在飞行器飞行时也有利于减小空气阻力，提高多个旋翼调控飞行姿态时的反应灵敏度。
当然在其他实施例中，旋翼设置的数量和排布方式不局限于以上所述或图中所示，例如也可以参照实施例7-10中设置旋翼的数量和排布方式。
上面结合附图对本发明的实施方式作了详细说明，但是本发明并不限于上述实施方式。即使对本发明做出各种变化，倘若这些变化属于本发明权利要求及其等同技术的范围之内，则仍落入在本发明的保护范围之中。

Claims (16)

  1. 一种结合多轴旋翼的自旋翼飞行器，其特征在于，包括
    机身；
    自旋翼，设置于所述机身的顶部；
    推进装置，设置于机身的尾部，其包括推进桨和驱动所述推进桨转动的推进驱动装置；
    至少一组多轴旋翼系统，用于为所述机身提供垂直上升的升力；每组所述多轴旋翼系统包括两对称设置于所述机身两侧的旋翼模块；每个所述旋翼模块包括至少一个旋翼，每个所述旋翼对应连接有旋翼驱动装置，所述旋翼模块通过支架安装于所述机身上；
    起落装置，设置于所述机身底部或所述支架上。
  2. 根据权利要求1所述的结合多轴旋翼的自旋翼飞行器，其特征在于，还包括：飞行动力系统，其包括：动力控制模块、燃油供应装置、电力供应装置；
    所述动力控制模块分别与所述燃油供应装置、所述推进驱动装置、所述旋翼驱动装置以及所述电力供应装置连接；
    根据所选的飞行状态，在垂直起降时控制所述电力供应装置给所述旋翼驱动装置提供电力，驱动所述旋翼转动；在平行飞行时控制燃油供应装置给所述推进驱动装置提供动力，或者控制所述电力供应装置给所述推进驱动装置提供电力，驱动所述推进桨转动。
  3. 根据权利要求1所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述多轴旋翼系统可拆卸式安装于所述机身；
    当安装所述多轴旋翼系统时，构成多轴旋翼自旋翼结合的飞行器；当拆卸所述多轴旋翼系统时，构成自旋翼飞行器。
  4. 根据权利要求2所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述支架包括旋翼架、支撑架，所述旋翼模块通过所述旋翼架安装于所述支撑架上；所述电力供应装置挂载于所述支撑架或所述机身外部。
  5. 根据权利要求4所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述多轴旋翼系统包括至少两个旋翼模块，所述旋翼模块关于所述机身的平均重心对称设置。
  6. 根据权利要求1所述的结合多轴旋翼的自旋翼飞行器，其特征在于，每个所述旋翼模块包括两个旋翼，且两所述旋翼上下同轴设置。
  7. 根据权利要求4所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述旋翼模块呈八字形对称分布于所述机身两侧。
  8. 根据权利要求1或2所述的结合多轴旋翼的自旋翼飞行器，其特征在于，每个所述旋翼配备的所述旋翼驱动装置为驱动电机，所述驱动电机设置于所述支架上。
  9. 根据权利要求8所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述推进驱动装置为燃油发动机或电力发动机，当推进驱动装置为燃油发动机时，通过所述燃油供应装置提供动力，当推进驱动装置为电力发动机时，通过所述电力供应装置提供电力。
  10. 根据权利要求1所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述支架采用可折叠和/或可伸缩的支架。
  11. 根据权利要求10所述的结合多轴旋翼的自旋翼飞行器，其特征在于，多个所述电力供应装置的重心与飞行器的平均重心位于同一竖向高度。
  12. 根据权利要求4所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述电力供应装置包括电池模块和电池舱，所述电池舱设置于所述支撑架下方，用于储放所述电池模块，且所述电池舱通过拆卸组件与所述支撑架可拆卸连接。
  13. 根据权利要求12所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述支撑架上与所述电池舱对应的位置处设置电池舱加强肋，所述电池舱通过拆卸组件吊挂于所述电池舱加强肋上；
    所述拆卸组件包括：
    连接座，其具有一插口，所述连接座固定于所述支撑架上；
    连接头，设置于所述电池舱上，所述连接头与所述插口匹配，所述连接头插设于所述插口内，实现所述连接头的径向限位；
    径向插件，沿所述连接座径向插设于所述连接座和所述连接头，实现所述连接头的轴向限位；
    固定件，用于限制所述径向插件径向移动脱离所述连接座和所述连接头。
  14. 根据权利要求13所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述拆卸组件还包括加强框架，所述电池舱和所述旋翼纵梁周侧均围设有所述加强框架；所述连接头设置于所述加强框架上。
  15. 根据权利要求4所述的结合多轴旋翼的自旋翼飞行器，其特征在于，所述支架还包括支撑梁，所述支撑梁一端与所述支撑架连接，一端与所述机身连接，所述支撑梁与所述机身、所述支撑架连接呈三角结构。
  16. 根据权利要求1所述的结合多轴旋翼的自旋翼飞行器，其特征在于，安装有所述多轴旋翼系统时，飞行器载人后的总重量小于或等于飞行器的最大许可起飞重量。
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