Distributed oil-driven multi-rotor aircraft and control method thereof

Abstract

The invention relates to a distributed oil-driven multi-rotor aircraft and a control method thereof.The aircraft comprises a transverse arm and a longitudinal arm which are arranged in a layered mode, the two ends of the transverse arm are each provided with a first rotor, and each first rotor is connected with a transverse engine through a transverse transmission structure; the two groups of transverse transmission structures and the two groups of transverse engines are arranged in the transverse arm; second rotors are arranged at the two ends of the longitudinal machine arm correspondingly, each second rotor is connected with a longitudinal engine through a longitudinal transmission structure, and the two longitudinal transmission structures and the two longitudinal engines are all arranged in the longitudinal machine arm. According to the multi-rotor aircraft, the fuel engine is adopted, distributed arrangement of a power system is achieved, long-endurance and long-voyage application of the aircraft can be achieved, the production cost of the aircraft can be reduced, and the use reliability of the multi-rotor aircraft can be improved; compared with a common helicopter, pitching and rolling control force is far away from the gravity center, and control is more agile and reliable.

Description

Technical Field

[0001] This invention belongs to the field of aircraft technology, specifically relating to a distributed gasoline-powered multirotor aircraft and its control method. Background Technology

[0002] With the development of vertical takeoff and landing (VTOL) aircraft technology, distributed-powered electric multirotor aircraft have become mainstream. These aircraft use multiple small electric power units combined to generate lift, achieving effective payload and flight control. This design reduces the design complexity and cost of VTOL aircraft, enabling their widespread application. However, due to the limited energy density of their energy storage batteries, distributed-powered electric multirotor aircraft generally suffer from short flight time and limited range. Furthermore, heavy-load electric VTOL aircraft, due to issues such as low energy density of their energy storage batteries, complex electric power control components, and large weight, exhibit significant differences in payload ratio, range, and flight time compared to traditional helicopters.

[0003] Traditional fuel-powered helicopters have significant advantages over electric vertical takeoff and landing aircraft in terms of payload capacity, range, and flight time. However, their pitch and roll control forces are close to the center of gravity, resulting in poor control agility and insufficient control reliability. Summary of the Invention

[0004] This invention relates to a distributed oil-powered multirotor aircraft and its control method, which can at least solve some of the defects of the prior art.

[0005] This invention relates to a distributed, gasoline-powered multirotor aircraft, comprising layered lateral and longitudinal arms.

[0006] The two ends of the lateral arm are respectively provided with first rotors. Each set of first rotors is connected to a set of lateral engines through a lateral transmission structure. The two sets of lateral transmission structures and the two sets of lateral engines are arranged in the lateral arm.

[0007] The longitudinal arm is equipped with a second rotor at each end. Each set of second rotors is connected to a set of longitudinal engines through a longitudinal transmission structure. Both sets of longitudinal transmission structures and both sets of longitudinal engines are arranged in the longitudinal arm.

[0008] As one implementation method, the lateral arm and the longitudinal arm are rotatably connected by a folding pivot, so that the aircraft has a folded state in which the lateral arm and the longitudinal arm are parallel, and an unfolded state in which the lateral arm and the longitudinal arm are perpendicular.

[0009] As one implementation method, two sets of transverse engines are symmetrically distributed with respect to the folding shaft, and two sets of transverse transmission structures are symmetrically distributed with respect to the folding shaft.

[0010] And / or, two sets of longitudinal engines are symmetrically distributed with respect to the folding shaft, and two sets of longitudinal transmission structures are symmetrically distributed with respect to the folding shaft.

[0011] As one implementation method, the rear end of the longitudinal arm is provided with a vertical tail.

[0012] As one implementation method, the vertical tail is a double-moving vertical tail.

[0013] As one implementation method, both sets of the lateral transmission structures include a lateral transmission shaft and a lateral belt clutch. The lateral belt clutch is connected to the corresponding lateral engine and the lateral transmission shaft, respectively. The lateral transmission shaft is connected to the corresponding first rotor.

[0014] And / or, both sets of longitudinal transmission structures include a longitudinal drive shaft and a longitudinal belt clutch, the longitudinal belt clutch being connected to the corresponding longitudinal engine and the longitudinal drive shaft, and the longitudinal drive shaft being connected to the corresponding second rotor.

[0015] The present invention also relates to a control method for the aforementioned distributed gasoline-powered multirotor aircraft, the control method comprising:

[0016] During takeoff and landing, vertical lift is generated by controlling the collective pitch of the two sets of first rotors and the two sets of second rotors, thus enabling the aircraft to take off and land vertically.

[0017] Specifically, pitch control of the aircraft is achieved by varying the collective pitch of the two sets of second rotors to cause differential changes in the lift of the two sets of second rotors; roll control of the aircraft is achieved by varying the collective pitch of the two sets of first rotors to cause differential changes in the lift of the two sets of first rotors.

[0018] As one implementation method, the control method further includes:

[0019] A vertical tail is provided at the rear end of the longitudinal arm;

[0020] During takeoff and landing, the yaw moment generated by the collective pitch differential of the two sets of first rotors and the two sets of second rotors is canceled out by the action of the vertical tail; and the yaw control of the aircraft during takeoff and landing is achieved by the yaw moment generated by the deflection of the vertical tail.

[0021] The present invention also relates to a control method for the aforementioned distributed gasoline-powered multirotor aircraft, the control method comprising:

[0022] During level flight, the pitch angle is generated by the collective pitch differential of the two sets of second rotors, and the lateral roll of the aircraft is controlled by the collective pitch differential of the two sets of first rotors.

[0023] The forward speed of the aircraft is changed by controlling the collective pitch of the two sets of first rotors and the two sets of second rotors, as well as by controlling the pitch of the aircraft.

[0024] As one implementation method, the control method further includes:

[0025] A vertical tail is provided at the rear end of the longitudinal arm;

[0026] During level flight, the yaw control of the aircraft is achieved by the yaw torque generated by the deflection of the vertical tail.

[0027] The present invention has at least the following beneficial effects:

[0028] This invention uses a fuel engine, enabling long endurance and long range applications for the aircraft; it employs two sets of lateral engines and two sets of lateral transmission structures to drive two sets of first rotors, and two sets of longitudinal engines and two sets of longitudinal transmission structures to drive two sets of second rotors, achieving a distributed power system layout. This not only reduces the production cost of the aircraft but also improves the reliability of the multi-rotor aircraft. The multiple engines also give the aircraft power redundancy.

[0029] The distributed nitro-powered multirotor aircraft provided by this invention, compared with ordinary helicopters, has pitch and roll control forces located far from the center of gravity, resulting in more agile and reliable control. Attached Figure Description

[0030] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0031] Figure 1 This is a schematic diagram of the structure of a distributed gasoline-powered multi-rotor aircraft provided in an embodiment of the present invention;

[0032] Figure 2 This is a diagram showing the equipment layout in the longitudinal arm;

[0033] Figure 3 This is a diagram showing the equipment layout in the horizontal arm;

[0034] Figure 4 This is a schematic diagram of a dual-axis bidirectional rudder provided in an embodiment of the present invention; Figure 5 This is a schematic diagram of another dual-axis bidirectional rudder provided in an embodiment of the present invention. Detailed Implementation

[0035] The technical solutions in the embodiments of the present invention will be clearly and completely described below. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0036] Example 1

[0037] like Figures 1-3 This invention provides a distributed gasoline-powered multirotor aircraft, comprising a layered horizontal arm 21 and a vertical arm 11.

[0038] The two ends of the lateral arm 21 are respectively provided with first rotors 22. Each set of first rotors 22 is connected to a set of lateral engines 131 through a lateral transmission structure. The two sets of lateral transmission structures and the two sets of lateral engines 131 are arranged in the lateral arm 21.

[0039] The longitudinal arm 11 is provided with a second rotor 12 at each end. Each set of second rotors 12 is connected to a set of longitudinal engines 231 through a longitudinal transmission structure. Both sets of longitudinal transmission structures and both sets of longitudinal engines 231 are arranged in the longitudinal arm 11.

[0040] In the aforementioned aircraft, the lateral arm 21 can be located above or below the longitudinal arm 11. Preferably, as shown... Figures 1-3 The lateral arm 21 is located above the longitudinal arm 11. When the aircraft provided in this embodiment is a helicopter, the fuselage can be mounted on the longitudinal arm 11.

[0041] Generally, the longitudinal arm 11 is parallel to the forward direction of the aircraft, while the lateral arm 21 is generally perpendicular to the longitudinal arm 11.

[0042] Among them, the horizontal arm 21 and the vertical arm 11 are both integral arms with complete load-bearing structures, which is conducive to bearing higher loads and can also improve the reliability of multi-rotor aircraft.

[0043] The use of an integral arm facilitates the arrangement of the transverse engine 131, transverse transmission structure, longitudinal engine 231, and longitudinal transmission structure, ensuring the safety of these devices. For example, the drive shafts in the transverse transmission structure and the longitudinal transmission structure can be set along the transverse arm 21 and the longitudinal arm 11, respectively.

[0044] Preferably, such as Figures 1-3The lateral arm 21 is a truss-type arm, which enables the lateral arm 21 to have a high strength-to-weight ratio and load capacity, as well as to ensure that the lateral arm 21 has excellent bending and torsional stiffness, effectively improving the flight stability and flight control accuracy of the aircraft, and is especially suitable for large multi-rotor aircraft.

[0045] Furthermore, the use of a truss-type arm for the lateral arm 21 facilitates the arrangement of the lateral engine 131 and the lateral transmission structure, etc.

[0046] Similarly, as Figures 1-3 The longitudinal arm 11 is preferably a truss arm.

[0047] In one embodiment, the lateral arm 21 and the longitudinal arm 11 are rotatably connected via a folding pivot, such that the aircraft has a folded state where the lateral arm 21 is parallel to the longitudinal arm 11 and an unfolded state where the lateral arm 21 is perpendicular to the longitudinal arm 11. Because an integral arm is used, in the folded state, there is no need to disassemble the lateral engine 131, lateral transmission structure, longitudinal engine 231, and longitudinal transmission structure—these power units. This avoids the disassembly of the load-bearing structures and these power units during transportation and the reassembly after transportation, thus greatly improving the ease of use of the multirotor aircraft. Simultaneously, it avoids the impact of frequent disassembly and reassembly of the load-bearing structures and power units on the aircraft's structural performance, thereby improving the reliability of the multirotor aircraft.

[0048] Generally, in the deployed state, it is necessary to ensure the positional stability of the transverse arm 21 and the longitudinal arm 11, that is, to lock the deployed state. Accordingly, a locking unit can be provided on the transverse arm 21 to lock the longitudinal arm 11 in the deployed state, and / or a locking unit can be provided on the longitudinal arm 11 to lock the transverse arm 21 in the deployed state. The above-mentioned locking units include, but are not limited to, locking structures such as electric locking pins, the specific structures of which will not be described in detail here.

[0049] In this embodiment, a fuel engine is used, that is, the aforementioned lateral engine 131 and longitudinal engine 231 are fuel engines, which can realize the long endurance and long range of the aircraft. The fuel tanks can be installed on the fuselage and connected to the two sets of lateral engines 131 and the two sets of longitudinal engines 231 through a distributed fuel line; alternatively, multiple fuel tanks can be distributed on the lateral arm 21 and the longitudinal arm 11, respectively supplying fuel to the two sets of lateral engines 131 and the two sets of longitudinal engines 231.

[0050] In addition, in this embodiment, two sets of lateral engines 131 and two sets of lateral transmission structures are used to drive two sets of first rotors 22, and two sets of longitudinal engines 231 and two sets of longitudinal transmission structures are used to drive two sets of second rotors 12, so as to realize the distributed arrangement of the power system, which can reduce the production cost of the aircraft and improve the reliability of the multi-rotor aircraft.

[0051] In one embodiment, both sets of lateral transmission structures include a lateral transmission shaft 133 and a lateral belt transmission assembly 132. The lateral belt transmission assembly 132 is connected to the corresponding lateral engine 131 and the lateral transmission shaft 133, respectively. The lateral transmission shaft 133 is connected to the corresponding first rotor 22. Furthermore, the lateral belt transmission assembly 132 can integrate a clutch to form a belt clutch.

[0052] Preferably, the two transverse drive shafts 133 are coaxially connected, for example, they can be connected by a coupling, or they can be connected as an integrated transverse output shaft, with both sets of transverse belt drive assemblies 132 connected to this integrated transverse output shaft. Based on the above structural design, the transverse power system can have power redundancy capability. In the event of failure of either transverse engine 131, the other transverse engine 131 can operate at reduced power, keeping the lift and torque of the aircraft balanced, ensuring the aircraft is controllable and flyable, thereby greatly improving the flight safety of the aircraft.

[0053] In one embodiment, both sets of longitudinal transmission structures include a longitudinal transmission shaft 233 and a longitudinal belt transmission assembly 232. The longitudinal belt transmission assembly 232 is connected to the corresponding longitudinal engine 231 and the longitudinal transmission shaft 233, respectively. The longitudinal transmission shaft 233 is connected to the corresponding second rotor 12. Furthermore, the longitudinal belt transmission assembly 232 can integrate a clutch to form a belt clutch.

[0054] Preferably, the two longitudinal drive shafts 233 are coaxially connected, for example, they can be connected by a coupling, or they can be connected as an integrated longitudinal output shaft, with both sets of longitudinal belt drive assemblies 232 connected to this integrated longitudinal output shaft. Based on the above structural design, the longitudinal power system can have power redundancy. In the event of failure of either longitudinal engine 231, the other longitudinal engine 231 can operate at reduced power, keeping the lift and torque of the aircraft balanced, ensuring the aircraft is controllable and flyable, thereby greatly improving the flight safety of the aircraft.

[0055] Preferably, the two sets of lateral engines 131 are symmetrically distributed with respect to the folding pivot, and the two sets of lateral transmission structures are symmetrically distributed with respect to the folding pivot; and / or, the two sets of longitudinal engines 231 are symmetrically distributed with respect to the folding pivot, and the two sets of longitudinal transmission structures are symmetrically distributed with respect to the folding pivot. The above structures can achieve stable, controllable, safe, and efficient flight of the aircraft, including but not limited to improving hovering balance performance, torque balance performance, decoupling control reliability, and flight control efficiency.

[0056] In one embodiment, such as Figure 1 and Figure 2 The longitudinal arm 11 has a vertical tail 4 at its rear end. Preferably, a double vertical tail structure is adopted; in addition, it is preferable to set the vertical tail 4 as a double-moving vertical tail, which can deflect around the horizontal axis and also around the vertical axis, which can greatly improve the flexibility and reliability of the deflection control of the vertical tail 4.

[0057] Based on the aforementioned distributed gasoline-powered multirotor aircraft, this embodiment also provides a control method for the aircraft, including:

[0058] (1) During the take-off and landing phase, the vertical lift is changed by controlling the collective pitch of the two sets of first rotors 22 and the two sets of second rotors 12, so as to realize the vertical take-off and landing of the aircraft;

[0059] The collective pitch of the two sets of second rotors 12 can change the difference in lift between the two sets of second rotors 12, enabling pitch control of the aircraft. When the lift of the front second rotor 12 increases and the lift of the rear second rotor 12 decreases, the aircraft generates a pitching moment; when the lift of the front second rotor 12 decreases and the lift of the rear second rotor 12 increases, the aircraft generates a pitching moment.

[0060] The collective pitch of the two sets of first rotors 22 changes, which causes a differential change in the lift of the two sets of first rotors 22. When the lift of the right first rotor 22 increases and the lift of the left first rotor 22 decreases, the aircraft generates a leftward roll moment; when the lift of the right first rotor 22 decreases and the lift of the left first rotor 22 increases, the aircraft generates a rightward roll moment.

[0061] The yaw moment generated by the collective pitch differential of the two sets of first rotors 22 and the two sets of second rotors 12 can be canceled out by the movement of the vertical tail 4. In addition, the yaw control of the aircraft during takeoff and landing can be achieved by the yaw moment generated by the deflection of the vertical tail 4.

[0062] (2) During the level flight phase, the pitch angle of the aircraft is generated by the collective pitch differential of the two sets of second rotors 12, and the lateral roll of the aircraft is controlled by the collective pitch differential of the two sets of first rotors 22.

[0063] The collective pitch of the two sets of second rotors 12 can change the difference in lift between the two sets of second rotors 12, enabling pitch control of the aircraft. When the lift of the front second rotor 12 increases and the lift of the rear second rotor 12 decreases, the aircraft generates a pitching moment; when the lift of the front second rotor 12 decreases and the lift of the rear second rotor 12 increases, the aircraft generates a pitching moment.

[0064] The collective pitch of the two sets of first rotors 22 can cause a differential change in the lift of the two sets of first rotors 22. When the lift of the right first rotor 22 increases and the lift of the left first rotor 22 decreases, the aircraft generates a roll moment to the left; when the lift of the right first rotor 22 decreases and the lift of the left first rotor 22 increases, the aircraft generates a roll moment to the right.

[0065] In addition, the forward speed of the aircraft can be changed by controlling the collective pitch of the two sets of first rotors 22 and the two sets of second rotors 12, as well as by controlling the pitch of the aircraft. Specifically, during cruise, the aircraft reduces its pitch angle, causing the two sets of first rotors 22 and the two sets of second rotors 12 to generate forward thrust. By controlling the collective pitch of the two sets of first rotors 22 and the two sets of second rotors 12, as well as by controlling the pitch angle of the aircraft, the forward thrust can be changed, thereby changing the forward speed of the aircraft.

[0066] In addition, by deflecting the vertical tail 4, the aircraft's forward heading can be kept stable.

[0067] The distributed nitro-powered multirotor aircraft provided in this embodiment, compared with ordinary helicopters, has pitch and roll control forces located far from the center of gravity, making control more agile and reliable.

[0068] Example 2

[0069] This invention provides a dual-axis bidirectional rudder, which can be used in the above embodiment 1 as the vertical tail 4.

[0070] When the above embodiment 1 adopts a dual vertical tail structure, preferably, both vertical tails 4 can adopt the dual-axis bidirectional rudder of this embodiment.

[0071] like Figure 4 The dual-axis bidirectional rudder includes a main rudder surface 411 and a secondary rudder surface 412. The main rudder surface 411 is connected to a main servo motor for driving it to rotate around a first rotating shaft 413. The secondary rudder surface 412 is rotatably mounted on the main rudder surface 411 via a second rotating shaft 414. The second rotating shaft 414 is perpendicular to the first rotating shaft 413 and parallel to the main rudder surface 411. The main rudder surface 411 is provided with a secondary servo motor for driving the secondary rudder surface 412 to rotate around the second rotating shaft 414.

[0072] The aforementioned main servo motor can be mounted on the longitudinal arm in Embodiment 1. The first rotating shaft 413 is connected to the output end of the main servo motor, and the main control surface 411 is connected to the first rotating shaft 413. The main servo motor can drive the main control surface 411 to rotate around the first rotating shaft 413, and it can also lock the first rotating shaft 413, thus keeping the main control surface 411 stationary and making it a stable surface.

[0073] In one embodiment, the first rotating shaft 413 is provided with a bearing for mounting to a longitudinal arm, the first rotating shaft 413 rotatably engaging with the bearing, and the main servo is a first rotary servo coaxially mounted to the bearing and connected to the first rotating shaft 413.

[0074] The second rotating shaft 414 is connected to the output end of the secondary servo, and the secondary control surface 412 is connected to the second rotating shaft 414. The secondary servo can drive the secondary control surface 412 to rotate around the second rotating shaft 414, and it can also lock the second rotating shaft 414, thus keeping the secondary control surface 412 relatively stationary relative to the main control surface 411, i.e., rotating together with the main control surface 411. Optionally, the second rotating shaft 414 is rotatably connected to the main control surface 411, the secondary control surface 412 is fixedly connected to the second rotating shaft 414, and the secondary servo is a second rotating servo coaxially connected to the second rotating shaft 414.

[0075] The rudder provided in this embodiment uses a combination of a main control surface 411 and a secondary control surface 412 to form a dual-axis, two-way control capability. For example, the first rotating shaft 413 can be installed horizontally, and the main control surface 411 and the secondary control surface 412 can be locked on the same plane, rotating together around the first rotating shaft 413 (horizontal rotating shaft). This allows the downwash airflow generated by the rotor to generate a yaw torque for yaw control of the aircraft. If the first rotating shaft 413 (horizontal rotating shaft) and the main control surface 411 are locked in a vertical position (i.e., the main control surface 411 is parallel to the vertical), becoming a stabilizing surface, the secondary control surface 412 can deflect around the second rotating shaft 414 (i.e., the vertical rotating shaft / vertical axis), using the incoming airflow to generate a yaw torque for yaw control of the aircraft. Therefore, the dual-axis, two-way rudder provided in this embodiment allows the aircraft to obtain yaw torque through the rudder for yaw control during vertical takeoff and landing and level flight, effectively improving the applicability and reliability of the rudder.

[0076] Based on the combined structure of the main control surface 411 and the secondary control surface 412, the area ratio of the main control surface 411 and the secondary control surface 412 can be configured according to the rotor downwash airflow velocity and the horizontal flight airflow velocity, so that the rudder can maintain reasonable control effectiveness under different operating conditions. In one embodiment, the area ratio of the main control surface 411 to the secondary control surface 412 is in the range of 1:(0.5~1).

[0077] Preferably, the secondary control surface 412 and the main control surface 411 are located in the same plane, which facilitates yaw control.

[0078] In one embodiment, such as Figure 4 The secondary control surface 412 is located on the side of the primary control surface 411 away from the first rotating shaft 413. In this design, further, as... Figure 4 Along the axial direction of the second rotating shaft 414, both ends of the main control surface 411 and the secondary control surface 412 are flush. For example, when the second rotating shaft 414 is parallel to the vertical direction, the height of the main control surface 411 is the same as the height of the secondary control surface 412, and their top ends and bottom ends are flush. This structure facilitates yaw control.

[0079] In another embodiment, such as Figure 5 The main control surface 411 is an L-shaped control surface with a missing corner area, and the secondary control surface 412 is disposed in the missing corner area. The L-shaped main control surface 411 includes a first control surface segment 4111 and a second control surface segment 4112. The first control surface segment 4111 is connected to the first rotating shaft 413, and the second control surface segment 4112 is connected to the end of the first control surface segment 4111 that is away from the first rotating shaft 413.

[0080] In the aforementioned structure employing an L-shaped main control surface 411, optionally, the secondary control surface 412 can be rotatably connected to the second control surface segment 4112. This facilitates the arrangement of the secondary servo and the second rotating shaft 414. For example, the second rotating shaft 414 can be connected to the middle position of the secondary control surface 412 (e.g., the axis of the second rotating shaft 414 passes through the center point of the secondary control surface 412). This effectively reduces the force exerted on the secondary servo when the secondary control surface 412 deflects around the second rotating shaft 414. An automatic locking unit capable of assisting in locking the secondary control surface 412 can be provided on the first control surface segment 4111. In another embodiment, the secondary control surface 412 is rotatably connected to the first control surface segment 4111, and the second control surface segment 4112 is provided with an automatic locking unit capable of assisting in locking the secondary control surface 412.

[0081] The aforementioned automatic locking unit is mainly used to assist in locking the secondary control surface 412 and the main control surface 411. This ensures the positional stability of the secondary control surface 412 relative to the main control surface 411 and reduces the force on the secondary servo. The automatic locking unit includes, but is not limited to, electric locking devices such as electric pins. Accordingly, a locking part (such as a pin seat that cooperates with the electric pin) is provided on the secondary control surface 412.

[0082] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A distributed, gasoline-powered multirotor aircraft, characterized in that, Including horizontal and vertical booms arranged in layers, The two ends of the lateral arm are respectively provided with first rotors. Each set of first rotors is connected to a set of lateral engines through a lateral transmission structure. The two sets of lateral transmission structures and the two sets of lateral engines are arranged in the lateral arm. The longitudinal arm is equipped with a second rotor at each end. Each set of second rotors is connected to a set of longitudinal engines through a longitudinal transmission structure. Both sets of longitudinal transmission structures and both sets of longitudinal engines are arranged in the longitudinal arm.

2. The distributed gasoline-powered multirotor aircraft as described in claim 1, characterized in that: The lateral arm and the longitudinal arm are rotatably connected by a folding pivot, so that the aircraft has a folded state in which the lateral arm and the longitudinal arm are parallel, and an unfolded state in which the lateral arm and the longitudinal arm are perpendicular.

3. The distributed gasoline-powered multi-rotor aircraft as described in claim 2, characterized in that: Two sets of lateral engines are symmetrically distributed with respect to the folding shaft, and two sets of lateral transmission structures are symmetrically distributed with respect to the folding shaft; And / or, two sets of longitudinal engines are symmetrically distributed with respect to the folding shaft, and two sets of longitudinal transmission structures are symmetrically distributed with respect to the folding shaft.

4. The distributed gasoline-powered multi-rotor aircraft as described in claim 1, characterized in that: The longitudinal arm has a vertical tail at its rear end.

5. The distributed gasoline-powered multi-rotor aircraft as described in claim 4, characterized in that: The vertical tail is a double-moving vertical tail.

6. The distributed gasoline-powered multi-rotor aircraft as described in claim 1, characterized in that: Both sets of lateral transmission structures include a lateral transmission shaft and a lateral belt clutch. The lateral belt clutch is connected to the corresponding lateral engine and the lateral transmission shaft, respectively. The lateral transmission shaft is connected to the corresponding first rotor. And / or, both sets of longitudinal transmission structures include a longitudinal drive shaft and a longitudinal belt clutch, the longitudinal belt clutch being connected to the corresponding longitudinal engine and the longitudinal drive shaft, and the longitudinal drive shaft being connected to the corresponding second rotor.

7. The control method for a distributed gasoline-powered multirotor aircraft as described in any one of claims 1 to 6, characterized in that, The control method includes: During takeoff and landing, vertical lift is generated by controlling the collective pitch of the two sets of first rotors and the two sets of second rotors, thus enabling the aircraft to take off and land vertically. Specifically, pitch control of the aircraft is achieved by varying the collective pitch of the two sets of second rotors to cause differential changes in the lift of the two sets of second rotors; roll control of the aircraft is achieved by varying the collective pitch of the two sets of first rotors to cause differential changes in the lift of the two sets of first rotors.

8. The control method as described in claim 7, characterized in that, The control method further includes: A vertical tail is provided at the rear end of the longitudinal arm; During takeoff and landing, the yaw moment generated by the collective pitch differential of the two sets of first rotors and the two sets of second rotors is canceled out by the action of the vertical tail; and the yaw control of the aircraft during takeoff and landing is achieved by the yaw moment generated by the deflection of the vertical tail.

9. The control method for a distributed gasoline-powered multirotor aircraft as described in any one of claims 1 to 6, characterized in that, The control method includes: During level flight, the pitch angle is generated by the collective pitch differential of the two sets of second rotors, and the lateral roll of the aircraft is controlled by the collective pitch differential of the two sets of first rotors. The forward speed of the aircraft is changed by controlling the collective pitch of the two sets of first rotors and the two sets of second rotors, as well as by controlling the pitch of the aircraft.

10. The control method as described in claim 9, characterized in that, The control method further includes: A vertical tail is provided at the rear end of the longitudinal arm; During level flight, the yaw control of the aircraft is achieved by the yaw torque generated by the deflection of the vertical tail.

Images