CN104859859A - Pneumatic optimization oil-electricity hybrid multi-rotor aircraft - Google Patents

Abstract

The invention provides an aerodynamically optimized fuel-electric hybrid multi-rotor aircraft, comprising a fuselage with upper and lower surfaces and a first cavity; a landing gear fixed on the lower surface of the fuselage; a fuel engine; semi-closed and fixed on the lower surface of the fuselage An installation platform for installing a fuel engine, which is formed with a second cavity connected to the first cavity; driven by the fuel engine to rotate to realize the flight of the aircraft and arranged on the output shaft of the fuel engine and located in the first or second cavity The power rotor; the electric rotor mechanism for receiving the remote first control signal to adjust the flight attitude of the aircraft; and the aerodynamic optimization mechanism for receiving the remote second control signal to eliminate the anti-twist when the aircraft flies and is arranged on the outer wall of the installation platform . The implementation of the present invention uses a fuel engine and an electric rotor, which not only has the characteristics of simple structure, stable flight, and easy operation, but also greatly improves the load, endurance time, cruising speed, size, and structure.

Description

An aerodynamically optimized fuel-electric hybrid multi-rotor aircraft

technical field

The invention relates to the field of multi-rotor aircraft, in particular to an aerodynamically optimized fuel-electric hybrid multi-rotor aircraft.

Background technique

Multirotor is an aircraft that can take off and land vertically. It has at least three rotor shafts on its fuselage, and each rotor shaft is equipped with an electric motor and a rotor that is driven by the aforementioned motor to rotate and form thrust. Various flight maneuvers are realized by changing the rotational speed between different rotors. Because the multi-rotor aircraft has the characteristics of simple structure, stable flight, easy operation, convenient portability, and low safety hazard, it is widely used in various fields at home and abroad.

In the prior art, multi-rotor aircrafts include electric multi-rotor aircrafts and hybrid multi-rotor aircrafts. Compared with traditional helicopters, the existing electric multi-rotor aircraft has been developed relatively maturely. Although it has a simple structure, the total distance between the symmetrical rotors is fixed, and the blades of each rotor are relatively short, and the linear velocity at the end of the blades is slow. , so that when a collision occurs, the impact force is small, it is not easy to be damaged, and it has the advantages of safety, but it is difficult to make it bigger. The main reason is that most of the existing electric multi-rotor aircraft use lithium polymer batteries as the power source. Because the energy density of power energy is much lower than that of biofuels, the battery life of electric multi-rotor aircraft is short, especially after the electric multi-rotor aircraft reaches a certain scale, the ratio of battery weight to take-off weight will increase significantly, resulting in a series of Difficult problems such as increased payload, increased blanking time, etc.

Although the existing hybrid multi-rotor aircraft can solve the above-mentioned problems that the electric multi-rotor aircraft occurs, when the fuel engine on the hybrid multi-rotor aircraft drives the power rotor to rotate, a relatively large anti-torsion will be formed. Theoretically speaking, the hybrid multicopter can overcome the anti-torsion and maintain stability by changing the speed difference of the brushless DC motor on it, but because the size difference between the power rotor and the attitude adjustment rotor on it is large Large, the generated air reaction force is not on the same order of magnitude, so it is difficult to achieve the purpose of overcoming anti-twist in practical applications.

The aforementioned two existing multi-rotor aircraft generally adopt a symmetrical structure in design, so that the existing multi-rotor aircraft has basically the same flight performance in all directions. Although it has high rapid maneuverability, the overall aerodynamic performance is not Optimal, and the lift provided is limited, which greatly increases the resistance during cruising, which is not conducive to improving the endurance time and cruising speed, especially in windy conditions, and lacks high wind resistance.

Contents of the invention

The technical problem to be solved by the embodiments of the present invention is to provide an aerodynamically optimized fuel-electric hybrid multi-rotor aircraft, which uses a fuel engine and an electric rotor, which not only has the characteristics of simple structure, stable flight, and easy operation, but also has the advantages of load, endurance, Cruising speed, size and structure have been greatly improved.

In order to solve the above technical problems, an embodiment of the present invention provides an aerodynamically optimized fuel-electric hybrid multi-rotor aircraft, the multi-rotor aircraft includes:

A fuselage, the fuselage comprising an upper surface, a lower surface, and a first cavity passing through the upper surface and the lower surface;

at least one landing gear, the landing gear is fixed to the lower surface of the fuselage;

a fuel engine;

A semi-enclosed installation platform for installing the fuel engine, the installation platform is fixed to the lower surface of the fuselage, and is formed with an opening facing the first cavity and communicating with the first cavity the second cavity;

A power rotor driven by the fuel engine to rotate and generate thrust to realize the flight of the multi-rotor aircraft. The power rotor is arranged on the output shaft of the fuel engine and is located in the second cavity or the first cavity. inside the cavity;

At least one electric rotor mechanism for receiving a remote first control signal to adjust the flight attitude of the multi-rotor aircraft; and

At least one aerodynamic optimization mechanism is used to receive the remote second control signal to eliminate the anti-twist of the multi-rotor aircraft during flight, and the aerodynamic optimization mechanism is arranged on the outer wall of the installation platform.

Wherein, the aerodynamic optimization mechanism includes a hinge, a slipstream rudder, a connecting rod and a slipstream steering gear; wherein,

There is at least one hinge, one end of each hinge is fixed on the outer wall of the installation platform, and the other end is fixed to the first surface of the slipstream rudder;

The slipstream rudder is rotatably mounted on the outer wall of the installation platform through the hinge, and its second surface is fixed to one end of the connecting rod;

The slipstream steering gear is fixed on the outer wall of the installation platform, which includes a rotating rocker connected to the other end of the connecting rod;

When the slipstream steering gear drives the rotating rocker arm to rotate, the connecting rod is used to control the slipstream steering gear to rotate toward or away from the slipstream steering gear around the hinge.

Wherein, the aerodynamic optimization mechanism further includes a fixed rocker arm, one end of the fixed rocker arm is fixed to the second surface of the slipstream rudder, and the other end is fixed to one end of the connecting rod.

Wherein, when the slipstream steering gear drives the rotating rocker arm to rotate clockwise, the slipstream steering gear is controlled to rotate around the hinge toward the slipstream steering gear through the connecting rod; When the steering gear drives the rotating rocker arm to rotate counterclockwise, the slipstream rudder is controlled to rotate around the hinge away from the slipstream steering gear through the connecting rod; or

When the slipstream steering gear drives the rotating rocker arm to rotate clockwise, the slipstream steering gear is controlled to rotate around the hinge away from the slipstream steering gear through the connecting rod; when the slipstream steering gear When the rotating rocker arm is driven to rotate counterclockwise, the slipstream rudder is controlled to rotate around the hinge toward the slipstream steering gear through the connecting rod.

Wherein, the multi-rotor aircraft also includes a plurality of platform support rods, one end of each platform support rod is fixed to the lower surface of the fuselage, and the other end is fixed to the outer wall of the installation platform, and the A slipstream rudder is provided on the outer wall of each platform support rod, and is connected to a corresponding slipstream rudder through the hinge.

Wherein, the electric rotor mechanism includes a brushless DC motor and an attitude adjustment rotor installed on the output shaft of the brushless DC motor.

Wherein, the multi-rotor aircraft also includes two support frames, and the fuselage also includes a first side and a second side; wherein,

The two supporting frames are respectively fixed on the first side and the second side and parallel to each other, and each supporting frame extends out of the machine body to both sides along the axial direction of the first side of the fuselage. Two installation positions for installing the brushless DC motor are formed outside the body.

Wherein, each support frame is fixed to the fuselage by screws.

Wherein, the fuselage further includes a third side; wherein, the thickness of the fuselage gradually decreases from the third side along the axis of the first side moving away from the third side, and the thickness of the fuselage Both the upper surface and the lower surface are formed as curved surfaces with a certain radian, and the curved surface radian of the upper surface is larger than that of the lower surface.

Wherein, the output shaft of the fuel engine is perpendicular to the central axis of the fuselage.

Implementing the embodiment of the present invention has the following beneficial effects:

1. In the embodiment of the present invention, since the multi-rotor aircraft uses the fuel engine to provide the main flight power, the mounting capacity and endurance time of the multi-rotor aircraft are greatly increased;

2. In the embodiment of the present invention, since the multi-rotor aircraft adopts the slipstream rudder design on the basis of the fuel engine, the anti-twist problem generated when the fuel engine is working is solved, so it has simple control and little impact on the flight performance of the multi-rotor aircraft specialty;

3. In the embodiment of the present invention, since the multi-rotor aircraft adopts the design of the installation angle difference between the fuselage and the rotor plane, that is, the design of the fuselage has a thickness difference and the upper and lower surfaces are curved surfaces with different radians, so that the multi-rotor aircraft During the forward flight, the fuselage is in a horizontal posture, which has small aerodynamic resistance during cruising, and at the same time, the fuselage generates a certain amount of aerodynamic lift, which can effectively improve the overall endurance time and cruising speed of the aircraft;

4. In the embodiment of the present invention, since the multi-rotor aircraft adopts an oil-electric hybrid design, when the fuel engine fails, the electric rotor mechanism can still be used to control the multi-rotor aircraft to make a safe emergency landing.

Description of drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings that need to be used in the description of the embodiments or the prior art. Obviously, the accompanying drawings in the following description are only These are some embodiments of the present invention. For those of ordinary skill in the art, obtaining other drawings based on these drawings still belongs to the scope of the present invention without any creative effort.

Fig. 1 is a schematic diagram of a three-dimensional structure of an aerodynamically optimized oil-electric hybrid multi-rotor aircraft provided by an embodiment of the present invention;

Fig. 2 is a top view three-dimensional structural schematic diagram of Fig. 1;

Fig. 3 is a partial enlarged view of point A in Fig. 1;

Fig. 4 is a partial enlarged view of point B in Fig. 3;

Fig. 5 is the partial plane structure schematic diagram of Fig. 2;

Fig. 6 is a schematic diagram of an application scenario of an aerodynamically optimized oil-electric hybrid multi-rotor aircraft provided by an embodiment of the present invention;

Fig. 7 is a schematic diagram of another application scenario of the aerodynamically optimized oil-electric hybrid multi-rotor aircraft provided by the embodiment of the present invention;

In the figure: 1-body, 11-upper surface, 12-lower surface, 13-first cavity, 14-first side, 15-second side, 16-third side, 2-landing gear, 3- Fuel engine, 4-installation platform, 41-second cavity, 5-power rotor, 6-electric rotor mechanism, 61-brushless DC motor, 62-attitude adjustment rotor, 7-aerodynamic optimization mechanism, 71-hinge, 72-slipstream rudder, 721-first side, 722-second side, 73-connecting rod, 74-slipstream steering gear, 741-rotating rocker arm, 75-fixed rocker arm, 8-platform support rod, 9- Support frame, 91-the installation position of the brushless DC motor.

Detailed ways

In order to make the object, technical solution and advantages of the present invention clearer, the present invention will be further described in detail below in conjunction with the accompanying drawings.

As shown in Figures 1 to 5, it is an aerodynamically optimized fuel-electric hybrid multi-rotor aircraft proposed in the embodiment of the present invention. The multi-rotor aircraft includes:

A fuselage 1, the fuselage 1 includes an upper surface 11, a lower surface 12, and a first cavity 13 passing through the upper surface 11 and the lower surface 12; wherein, the fuselage 1 is made of carbon fiber material;

At least one landing gear 2, the landing gear 2 is fixed to the lower surface 12 of the fuselage 1;

a fuel engine 3;

A semi-enclosed installation platform 4 for installing the fuel engine 3, the installation platform 4 is fixed to the lower surface 12 of the fuselage 1, and it forms a second opening facing the first cavity 13 and communicating with the first cavity 13. cavity 41;

A power rotor 5 driven by the fuel engine 3 to rotate and generate thrust to realize the flight of the multi-rotor aircraft. The power rotor 5 is arranged on the output shaft of the fuel engine 3 and is located in the second cavity 41 or the first cavity 13;

At least one electric rotor mechanism 6 for receiving the remote first control signal to adjust the flight attitude of the multi-rotor aircraft; and

At least one aerodynamic optimization mechanism 7 is used to receive the remote second control signal to eliminate anti-twist when the multi-rotor aircraft is flying. The aerodynamic optimization mechanism 7 is arranged on the outer wall of the installation platform 4 .

It should be noted that the fuselage 1 should also be provided with a flight control system, a fuel tank, a throttle servo, a power supply unit, a data transceiver unit for receiving remote control signals, and various sensors for the multi-rotor aircraft to fly. The first control signal and the second control signal will be sent to the flight control system through the data transceiver unit, so that the flight control system outputs different control commands according to different control signals, including the start or close command of the fuel engine 3, the speed of the electric rotor mechanism 6 Adjustment instructions, instructions such as starting and closing of the pneumatic optimization mechanism 7 and changes in the direction of torque.

It should be noted that the aerodynamic optimization mechanism 7 will be controlled according to the rotation direction of the power rotor 5, so that the aerodynamic optimization mechanism 7 forms a torque opposite to that of the power rotor 5, thereby achieving the purpose of eliminating anti-twist, such as when the power rotor 5 is counterclockwise When the direction is rotated, the pneumatic optimization mechanism 7 is controlled to form a counterclockwise torque.

It can be understood that the output shaft of the fuel engine 3 should be perpendicular to the central axis of the fuselage 1, so that the power rotor 5 is parallel to the central axis of the fuselage 1, thereby generating aerodynamic lift. In order to reduce air resistance, the semi-closed installation platform 4 is adopted, which can effectively improve the overall flight time and cruising speed of the aircraft.

In the embodiment of the present invention, the pneumatic optimization mechanism 7 includes a hinge 71, a slipstream rudder 72, a connecting rod 73, and a slipstream rudder 74; wherein,

There is at least one hinge 71, one end of each hinge 71 is fixed on the outer wall of the installation platform 4, and the other end is fixed to the first surface 721 of the slipstream rudder 72;

The slipstream rudder 72 is rotatably mounted on the outer wall of the installation platform 4 through the hinge 71, and its second surface 722 is fixed to one end of the connecting rod 73;

The slipstream steering gear 74 is fixed on the outer wall of the installation platform 4, and it includes a rotating rocker 741 connected to the other end of the connecting rod 73;

When the slipstream steering gear 74 drives the rotating rocker arm 741 to rotate, the connecting rod 73 controls the slipstream steering gear 72 to rotate toward or away from the slipstream steering gear 74 around the hinge 71 .

Since the slipstream steering gear 74 drives the rotating rocker arm 741 to rotate in two directions, clockwise and counterclockwise, there are two corresponding rotation directions of the slipstream steering gear 72, as follows:

(1) When the slipstream steering gear 74 drives the rocker arm 741 to rotate clockwise, the connecting rod 73 controls the slipstream steering gear 72 to rotate around the hinge 71 toward the slipstream steering gear 74; when the slipstream steering gear 74 drives the rotating rocker arm When 741 rotates counterclockwise, the slipstream rudder 72 is controlled by the connecting rod 73 to rotate around the hinge 71 away from the slipstream servo 74;

(2) When the slipstream steering gear 74 drives the rotating rocker arm 741 to rotate clockwise, the connecting rod 73 controls the slipstream steering gear 72 to rotate around the hinge 71 away from the slipstream steering gear 74; when the slipstream steering gear 74 drives the rotating rocker arm When 741 rotates counterclockwise, the connecting rod 73 controls the slipstream rudder 72 to rotate around the hinge 71 toward the slipstream servo 74 .

In order to enhance the reliability of the connection between the second surface 722 of the slipstream rudder 72 and the connecting rod 73, the aerodynamic optimization mechanism 7 also includes a fixed rocker arm 75, an end of the fixed rocker arm 75 and the second surface 722 of the slipstream rudder 72. Phase is fixed, and the other end is fixed with one end of connecting rod 73.

In one embodiment, the hinge 71 , the slipstream rudder 72 , the connecting rod 73 , the slipstream steering gear 74 and the fixed rocker arm 75 are all fixed by glue.

In order to further improve the elimination of the torsion produced by the power rotor 5, the multi-rotor aircraft also includes a plurality of platform support rods 8, one end of each platform support rod 8 is fixed with the lower surface 12 of the fuselage 1, and the other end is connected with the installation The outer wall of the platform 4 is fixed, and a slipstream steering gear 74 is provided on the outer sidewall of each platform support rod 8, and is connected to a corresponding slipstream rudder 72 by a hinge 71, so that the slipstream rudder 72 is distributed in the On the support bar 8 of different platforms, the air resistance of the slipstream rudder 72 has been increased.

The embodiment of the present invention adopts a hybrid design of oil and electricity. When the fuel engine 3 fails, the electric rotor mechanism 6 can still control the multi-rotor aircraft to make a safe emergency landing. Therefore, the electric rotor mechanism 6 includes a brushless DC motor 61 and is installed in a brushless DC The attitude adjustment rotor 62 on the output shaft of the motor 61 makes it possible to control the brushless DC motor 61 to adjust the rotational speed of the attitude adjustment rotor 62 through the first control signal, so as to achieve the purpose of a safe forced landing of the multi-rotor aircraft.

Usually, the multi-rotor aircraft adopts two support frames 9 to install the electric rotor mechanism 6, and the two support frames 9 are respectively fixed on two opposite sides or end faces of the fuselage 2, and are designed in a symmetrical structure, so the multi-rotor aircraft also includes Two support frames 9, the fuselage 1 also includes a first side 14 and a second side 15; wherein, the two support frames 9 are respectively fixed on the first side 14 and the second side 15 of the fuselage 1 and then parallel to each other, and Each support frame 9 extends out of the fuselage 1 to both sides along the axial direction of the first side 14 of the fuselage 1 to form two installation positions 91 for installing the brushless DC motor 61 . Wherein, each supporting frame 9 is fixed to the fuselage 1 by screws.

In the embodiment of the present invention, by introducing the aerodynamic design concept of traditional aircraft, the outline of the fuselage 1 can be designed to roughly present a semi-elliptical shape, and the wing shape design of a fixed-wing aircraft is adopted. At this time, the fuselage 1 also includes a third side 16; wherein, the thickness of the fuselage 1 decreases gradually when moving along the axis of the first side 14 away from the third side 16 from the third side 16, so that the fuselage has a thickness difference, and the upper surface 11 and the lower surface 12 of the fuselage 1 are both It is formed as a curved surface with a certain radian, and the curved surface radian of the upper surface 11 is greater than that of the lower surface 12 . When the fuselage 1 moves relative to the air, the distance traveled by the air flowing through the upper surface 11 at the same time is longer than the distance of the air flowing through the lower surface 12, so the relative speed of the air on the upper surface 11 is faster than that of the lower surface. The air at 12 is fast, and according to Panulli's theorem, it can be known that "the pressure generated by the fluid on the surrounding substances is inversely proportional to the relative velocity of the fluid", so the pressure generated by the upper surface 11 of the fuselage 1 is smaller than that generated by the lower surface 12 of the fuselage 1 Pressure, so that the fuselage 1 can generate a certain lift, and can ensure that the multi-rotor aircraft has a certain maneuverability in different directions, and can improve the accuracy of fixed-point hovering.

The working principle of the aerodynamically optimized oil-electric hybrid multi-rotor aircraft provided by the embodiment of the present invention is: when the power rotor 5 of the multi-rotor aircraft rotates counterclockwise (as shown in the direction of arrow a in Figure 5), the body of the multi-rotor aircraft will be generated clockwise. In order to overcome this torque, the flight control system sends instructions to the aerodynamic optimization mechanism 7 to make the slipstream steering gear 74 control the slipstream rudder 72 to deflect in the same direction (as shown in Fig. 5 middle arrow b movement direction), so that the air flow direction is deflected (as shown in the arrow b movement direction in Figure 5), and a corresponding thrust in the opposite direction (as shown in the arrow a movement direction of Figure 5) will be generated; at this time, through the airflow The reverse thrust produced reaches the purpose of counteracting the torsion force produced by the power rotor 5 of the multi-rotor aircraft;

During the take-off and landing process of the multi-rotor aircraft, as shown in Figure 6, the rotation speed of the attitude adjustment rotor 61 is adjusted to be consistent, so that the rotation plane of the attitude adjustment rotor 61 is parallel to the horizontal plane. At this time, the multi-rotor aircraft fuselage 1 maintains a certain angle of attack. Engine 3 drives the rotation of power rotor 5 to control multi-rotor aircraft to take off and land;

During the forward flight of the multi-rotor aircraft, as shown in Figure 7, the attitude adjustment rotor 61 rotating speed of the third side 16 away from the fuselage 1 is adjusted to increase, and at this moment, the lift produced by the attitude adjustment rotor 61 away from the third side 16 Greater than the lift generated by the attitude adjustment rotor 61 toward the third side 16, the multi-rotor aircraft will have a tendency to bow its head, and the multi-rotor aircraft will fly forward through the horizontal component of the lift generated by each attitude adjustment rotor 61. Since the angle of attack of the fuselage 1 of the multi-rotor aircraft is reduced, the horizontal aerodynamic resistance in the forward direction of the multi-rotor aircraft is reduced, so that it has optimal aerodynamic performance in the forward direction, and the fuselage 1 will also generate a certain lift to achieve Increased cruising speed of multi-rotor aircraft;

In the same way, corresponding to the backward or lateral flight process of the multi-rotor aircraft, the principle is similar to the forward flight process of the multi-rotor aircraft, and will not be repeated here.

Implementing the embodiment of the present invention has the following beneficial effects:

1. In the embodiment of the present invention, since the multi-rotor aircraft uses the fuel engine to provide the main flight power, the mounting capacity and endurance time of the multi-rotor aircraft are greatly increased;

2. In the embodiment of the present invention, since the multi-rotor aircraft adopts the slipstream rudder design on the basis of the fuel engine, the anti-twist problem generated when the fuel engine is working is solved, so it has simple control and little impact on the flight performance of the multi-rotor aircraft specialty;

3. In the embodiment of the present invention, since the multi-rotor aircraft adopts the design of the installation angle difference between the fuselage and the rotor plane, that is, the design of the fuselage has a thickness difference and the upper and lower surfaces are curved surfaces with different radians, so that the multi-rotor aircraft During the forward flight, the fuselage is in a horizontal posture, which has small aerodynamic resistance during cruising, and at the same time, the fuselage generates a certain amount of aerodynamic lift, which can effectively improve the overall endurance time and cruising speed of the aircraft;

4. In the embodiment of the present invention, since the multi-rotor aircraft adopts an oil-electric hybrid design, when the fuel engine fails, the electric rotor mechanism can still be used to control the multi-rotor aircraft to make a safe emergency landing.

The above disclosure is only a preferred embodiment of the present invention, which certainly cannot limit the scope of rights of the present invention. Therefore, equivalent changes made according to the claims of the present invention still fall within the scope of the present invention.

Claims (10)

1. An aerodynamic optimization oil-electric hybrid multi-rotor aircraft, is characterized in that, the multi-rotor aircraft comprises: A fuselage (1), the fuselage (1) comprising an upper surface (11), a lower surface (12) and a first cavity (13) passing through the upper surface (11) and the lower surface (12) ); At least one landing gear (2), the landing gear (2) is fixed to the lower surface (12) of the fuselage (1); a fuel engine (3); A semi-enclosed installation platform (4) for installing the fuel engine (3), the installation platform (4) is fixed to the lower surface (12) of the fuselage (1), and it is formed with an opening facing the A second cavity (41) communicating with the first cavity (13) and the first cavity (13); A power rotor (5) driven by the fuel engine (3) to rotate and generate thrust to realize the flight of the multi-rotor aircraft, the power rotor (5) is arranged on the output shaft of the fuel engine (3), and located in the second cavity (41) or the first cavity (13); At least one electric rotor mechanism (6) for receiving a remote first control signal to adjust the flight attitude of the multi-rotor aircraft; and At least one aerodynamic optimization mechanism (7) for receiving a remote second control signal to eliminate anti-twist of the multi-rotor aircraft during flight, the aerodynamic optimization mechanism (7) is arranged on the outer wall of the installation platform (4) . 2. The multi-rotor aircraft according to claim 1, characterized in that, the aerodynamic optimization mechanism (7) includes a hinge (71), a slipstream rudder (72), a connecting rod (73) and a slipstream steering gear ( 74); where, There is at least one hinge (71), one end of each hinge (71) is fixed on the outer wall of the installation platform (4), and the other end is connected to the first surface of the slipstream rudder (72). (721) phase fixed; The slipstream rudder (72) is rotatably installed on the outer wall of the installation platform (4) through the hinge (71), and its second surface (722) is fixed to one end of the connecting rod (73) ; The slipstream steering gear (74) is fixed on the outer wall of the installation platform (4), which includes a rotating rocker arm (741) connected to the other end of the connecting rod (73); When the slipstream steering gear (74) drives the rotating rocker arm (741) to rotate, the slipstream steering wheel (72) is controlled by the connecting rod (73) to move toward or away from the hinge (71) The slipstream steering gear (74) rotates. 3. The multi-rotor aircraft according to claim 2, characterized in that, the aerodynamic optimization mechanism (7) further comprises a fixed rocker arm (75), one end of the fixed rocker arm (75) and the slipstream rudder The second surface (722) of (72) is fixed, and the other end is fixed to one end of the connecting rod (73). 4. The multi-rotor aircraft according to claim 3, characterized in that, when the slipstream steering gear (74) drives the rotating rocker arm (741) to rotate clockwise, it is controlled by the connecting rod (73) The slipstream rudder (72) rotates around the hinge (71) toward the slipstream steering gear (74); when the slipstream steering gear (74) drives the rotating rocker arm (741) to rotate counterclockwise , control the slipstream rudder (72) to rotate around the hinge (71) away from the slipstream rudder (74) through the connecting rod (73); or When the slipstream steering gear (74) drives the rotating rocker arm (741) to rotate clockwise, the slipstream steering wheel (72) is controlled to deviate from the hinge (71) through the connecting rod (73) The slipstream steering gear (74) rotates; when the slipstream steering gear (74) drives the rotating rocker arm (741) to rotate counterclockwise, the slipstream steering gear ( 72) Rotate around the hinge (71) towards the slipstream steering gear (74). 5. The multi-rotor aircraft according to claim 4, characterized in that, the multi-rotor aircraft also includes a plurality of platform support rods (8), and one end of each platform support rod (8) is connected to the fuselage ( The lower surface (12) of 1) is fixed, and the other end is fixed to the outer wall of the installation platform (4), and the outer wall of each platform support rod (8) is provided with a slipstream steering gear (74), and are all connected with a corresponding slipstream rudder (72) through the hinge (71). 6. The multi-rotor aircraft according to claim 5, characterized in that, the electric rotor mechanism (6) includes a brushless DC motor (61) and an attitude motor mounted on the output shaft of the brushless DC motor (61). Adjust rotor (62). 7. The multi-rotor aircraft according to claim 6, characterized in that, the multi-rotor aircraft also includes two support frames (9), and the fuselage (1) also includes a first side (14) and a second side side (15); where, The two support frames (9) are respectively fixed on the first side (14) and the second side (15) and parallel to each other, and each support frame (9) is along the fuselage (1 ) The first side ( 14 ) axially extends out of the fuselage ( 1 ) on both sides to form two mounting positions ( 91 ) for mounting the brushless DC motor ( 61 ). 8. The multi-rotor aircraft according to claim 7, characterized in that, each support frame (9) is fixed to the fuselage (1) by screws. 9. The multi-rotor aircraft according to claim 8, characterized in that, the fuselage (1) also includes a third side (16); wherein, the thickness of the fuselage (1) is from the third side ( 16) From the beginning, the movement along the axis of the first side (14) away from the direction of the third side (16) gradually decreases, and the upper surface (11) and lower surface (12) of the fuselage (1) are formed as The curved surface has a certain curvature, and the curved surface curvature of the upper surface (11) is greater than the curved surface curvature of the lower surface (12). 10. The multi-rotor aircraft according to claim 9, characterized in that, the output shaft of the fuel engine (3) is perpendicular to the central axis of the fuselage (1).