CN115303479A

CN115303479A - Multi-rotor combined helicopter

Info

Publication number: CN115303479A
Application number: CN202211073301.8A
Authority: CN
Inventors: 熊子见; 徐元铭; 王舟阳
Original assignee: Beihang University
Current assignee: Beihang University
Priority date: 2022-09-02
Filing date: 2022-09-02
Publication date: 2022-11-08

Abstract

A multi-rotor combined helicopter comprises a fuselage frame, a main rotor, wings, side rotors, a front rotor and a wing tilting mechanism; the main rotor wing is arranged right above the fuselage frame; the wings are connected to the fuselage frame through wing tilting structures; the tail ends of the wings are correspondingly connected with the side rotor wings; the side rotor wing is used for providing tension and yaw control moment when flying ahead in a fixed wing mode; also for providing upward lift in helicopter mode; the front rotor is arranged at the head part of the fuselage frame and is used for providing pitching maneuvering moment and providing upward lift force in a helicopter mode or generating maneuvering moment by changing the rotating speed difference with the main rotor; the invention can realize the conversion of the helicopter mode and the fixed wing flight mode by using the same set of power system under the condition of not greatly changing the attitude of the airplane, and simultaneously has the vertical take-off and landing capability, high suspension efficiency, large takeoff load and long voyage and high cruising speed of the fixed wing aircraft under the condition of ensuring lower waste weight.

Description

Multi-rotor combined helicopter

Technical Field

The invention relates to the technical field of unmanned aerial vehicles, in particular to a multi-rotor combined helicopter.

Background

At present, the design of a high-speed helicopter is always a big problem in the field of helicopters, a pair or a plurality of pairs of large rotors are used for providing lift force for a traditional helicopter in hovering and forward flying states, shock waves can be generated on forward blades of the helicopter in a high-speed flat flying state, and a large counter flow area can be generated by stalling backward blades, so that the efficiency of the rotors can be damaged by the serious asymmetric lift force, and the further speed increase of the helicopter is limited.

When the helicopter flies, the rotary reaction torque of the main rotor can enable the helicopter to rotate in the direction opposite to the rotation direction of the main rotor, the traditional single-rotor helicopter offsets the reaction torque of the main rotor by installing the tail rotor, the tail rotor can consume about 15% of the engine power without generating lift force, energy waste is caused, and the tail rotor needs to be provided with a tail beam and other structural supports, so that the weight of the whole helicopter is further increased.

In the prior art, the attitude control is realized by installing a tilting disk and a variable pitch mechanism at the hubs of two pairs of rotors, such as a coaxial helicopter; the control mechanism is quite complex and has low reliability and insufficient control torque, so that the maneuvering capability of the aircraft is limited. And the resistance generated by the high hub of the aircraft during high-speed forward flight even exceeds 50% of the full-aircraft resistance, which is not beneficial to the improvement of the cruising speed of the aircraft.

The vertical takeoff state and the fixed wing forward flight state in the existing multi-rotor wing-fixed wing combined unmanned aerial vehicle need two sets of different power systems, so that the waste weight under two modes of a plurality of rotor wings and fixed wings is increased to some extent, and the helicopter is difficult to surpass a helicopter with the same size.

And the transition process when traditional rotor craft that verts carries out the mode and switches is overlength, leads to the aircraft to be in nonlinear control district for a long time, causes the controllability to worsen, and the accident is many.

Therefore, how to provide a multi-rotor compound helicopter which can realize multi-mode flight and improve the flight efficiency and maneuverability is a problem to be solved in the field.

Disclosure of Invention

In view of this, the invention provides a multi-rotor combined helicopter, which can realize the conversion between a helicopter mode and a fixed-wing flight mode, and has the high suspension efficiency, large takeoff load and long range and high speed of a fixed-wing aircraft of the helicopter.

In order to achieve the purpose, the invention adopts the following technical scheme:

a multi-rotor combined helicopter comprises a fuselage frame, a main rotor, wings, side rotors, a front rotor and a wing tilting mechanism;

the main rotor wing is arranged right above the airframe frame;

the wings are connected to the fuselage frame through the wing tilting structures;

the tail ends of the wings are correspondingly connected with the side rotor wings; the side rotor wing is used for providing tension and yaw control moment when the fixed wing flies forwards in a mode of a fixed wing; also for providing upward lift in helicopter mode;

the front rotor is mounted to the head of the fuselage frame for providing a pitching steering torque and providing upward lift in helicopter mode or by varying the difference in rotational speed from the main rotor to produce a steering torque.

Furthermore, the wing tilting mechanism comprises a steering engine, a rudder disc sleeve and a fixed base; the fixed base is provided with a clamping groove, the steering engine is fixed in the clamping groove, and the output ends of two sides of the steering engine are respectively and fixedly connected with the corresponding inner wall of the steering wheel sleeve;

the fixed base is fixed inside the machine body frame;

the root part of the wing is provided with a wing rotating shaft; the wing rotating shaft is fixedly connected into the rudder disk sleeve.

Further, the front rotor changes the tension direction of the main rotor by providing pitching operation torque when the modes are switched, and the mode transition efficiency is improved.

Furthermore, the side rotor is also used for providing a rolling control moment through asymmetric rotating speed adjustment, providing a yawing control moment in cooperation with asymmetric tilting of the wing, and generating the yawing control moment by increasing or reducing the rotating speed difference between the side rotor and the main rotor.

Further, the main rotor is also used to stall and automatically lock in a direction parallel to the fuselage to reduce drag after switching to fixed wing mode.

Furthermore, the front rotor and the side rotors on two sides of the fuselage frame are symmetrically distributed in a triangular mode by taking the center of mass of the fuselage frame as the center.

Further, the rotation direction of the front rotor and the side rotors is opposite to the rotation direction of the main rotor, and the front rotor and the side rotors are used for providing lift force and offsetting reaction torque of the main rotor when a helicopter is in a mode

Furthermore, canard wings are arranged on two sides of the head of the fuselage frame respectively and used for increasing lift force and moving the full-aircraft aerodynamic focus forwards when the aircraft works in a fixed wing mode.

Furthermore, a horizontal tail wing and a vertical tail wing are arranged at the tail part of the machine body frame, and the two ends of the horizontal tail wing are respectively and correspondingly connected with the vertical tail wing to form an H-shaped layout.

Furthermore, a front undercarriage and a rear undercarriage are respectively arranged on the front side and the rear side of the bottom surface of the fuselage frame. The invention has the beneficial effects that the top end of the fuselage frame is provided with the airspeed head which is used for measuring the airspeed of the aircraft in the fixed wing mode flight:

according to the technical scheme, compared with the prior art, the invention discloses and provides the multi-rotor wing combined helicopter, which can realize the conversion between the helicopter mode and the fixed wing flight mode, and has the advantages of high suspension efficiency, large takeoff load and long range and high speed of a fixed wing aircraft;

the invention has the vertical take-off and landing capability and hovering capability which are not possessed by the fixed-wing unmanned aerial vehicle, is not limited by places, can execute tasks without special airport runways, has good low-altitude and low-speed performance, can adapt to the urban low-altitude environment with more obstacles, and can realize high-speed flat flight in open suburbs after turning into the fixed-wing mode.

Drawings

In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the provided drawings without creative efforts.

FIG. 1 is a schematic view of a multi-rotor hybrid helicopter according to the present invention;

FIG. 2 is a schematic diagram of a wing tilting mechanism according to the present invention;

FIG. 3 is a schematic cross-sectional view of an exemplary task in an embodiment of the present invention;

FIG. 4 (a) is a front view of a fixed wing mode of a multi-rotor helicopter in accordance with the present invention;

FIG. 4 (b) is a schematic side view of a fixed wing mode of a multi-rotor helicopter in accordance with the present invention;

FIG. 4 (c) is a schematic top view of a fixed wing of a multi-rotor helicopter in accordance with the present invention;

FIG. 5 (a) is a schematic helicopter modal front view of a multi-rotor hybrid helicopter provided in accordance with the present invention;

FIG. 5 (b) is a schematic side view of a helicopter in a multi-rotor hybrid helicopter in accordance with the present invention;

FIG. 5 (c) is a schematic top view of a multi-rotor helicopter in accordance with the present invention;

FIG. 6 is a schematic view of a flight control method according to the present invention;

wherein, 1-airspeed head, 2-front rotor wing, 3-canard wing, 4-nose landing gear, 5-fuselage frame; 6-wing tilting mechanism, 7-wing, 8-left rotor wing, 9-rear landing gear and 10-vertical tail wing; 11-horizontal tail; 12-right rotor, 13-main rotor; 14-wing root, 15-fixed base, 16-wing rotating shaft, 17-rudder disk sleeve and 18-steering engine.

Detailed Description

The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be obtained by a person skilled in the art without making any creative effort based on the embodiments in the present invention, belong to the protection scope of the present invention.

The embodiment of the invention discloses a multi-rotor combined helicopter, which comprises a fuselage frame 5, a main rotor 13, wings 7, side rotors, a front rotor 2 and a wing tilting mechanism 6, wherein the main rotor is arranged on the fuselage frame;

the main rotor 13 is arranged right above the fuselage frame 5;

the wings 7 are connected to the fuselage frame 5 through wing tilting structures 6;

the tail end of the wing 7 is correspondingly connected with the side rotor; the side rotor wing is used for providing pulling force and yaw control moment when the fixed wing flies forwards in a mode; also for providing upward lift in helicopter mode;

the front rotor 2 is mounted on the head of the fuselage frame 5 for providing a pitching steering torque and, in helicopter mode, an upward lift or steering torque by varying the difference in rotational speed with the main rotor 13. Wherein, the front part of the fuselage frame 5 is dug and is equipped with the duct for install preceding rotor 2.

Wherein, the main rotor 13 is arranged right above the mass center of the airplane and is used for providing lift force when the helicopter is in a mode; the hub of the aircraft does not need to be provided with any additional control systems such as a periodic variable pitch tilting disk, a flapping hinge, a shimmy hinge, a variable pitch hinge and the like, the attitude control of the aircraft is completely realized by a front rotor wing, a side rotor wing and a tilting wing, the number of main rotor wing parts is small, and the reliability is high; main rotor 13 contains two paddles, can be by motor (or fuel engine) direct drive, and to unmanned aerial vehicle, its paddle can be for fixed pitch formula, and pulling force size control realizes through changing motor speed. For a medium-sized unmanned aerial vehicle, a blade of the medium-sized unmanned aerial vehicle can be provided with a lead-lag hinge system, and the periodic pitch variation is realized by controlling the acceleration and deceleration of a motor in one rotation period. For a large unmanned aerial vehicle, the root of a blade of the unmanned aerial vehicle can be provided with a variable pitch hinge, and the blade is connected with an automatic inclinator through a pull rod to realize periodic variable pitch control.

The wings 7 are used to generate lift in a flat flight state, and when the helicopter is in modal suspension, a part of the reaction torque of the main rotor 13 can be offset by means of the washing flow of the main rotor 13 in an asymmetric deflection mode.

The two side rotors can provide upward lift force in a helicopter mode, simultaneously can provide a rolling control moment through asymmetric adjusting rotating speed, can provide a yawing control moment by matching with asymmetric tilting of the wing 7, and can generate the yawing control moment by increasing or reducing the rotating speed difference between the two side rotors and the main rotor 13. The fixed wing provides tension when flying ahead in a mode, and simultaneously can provide yaw control moment by adjusting the rotating speed.

In another embodiment, the wing tilting mechanism 6 comprises a steering engine 18, a rudder disk sleeve 17 and a fixed base 15; a clamping groove is formed in the fixed base 15, the steering engine 18 is fixed in the clamping groove, and the output ends of two sides of the steering engine 18 are respectively and fixedly connected with the inner wall of the corresponding steering wheel sleeve 17; the fixed base 15 is fixed inside the machine body frame 5; the root of the wing 7 is provided with a wing rotating shaft 16; the wing shaft 16 is fixedly connected into the rudder disk sleeve 17.

In another embodiment, the front rotor 2 and the side rotors on both sides of the fuselage frame 5 are distributed in a triangular symmetry with the center of mass of the fuselage frame 5 as the center. The main rotor 13 does not need any tilting disk and a variable-pitch mechanism, the attitude control in the hovering state is completely completed by three small rotors, the complexity of the mechanism is simplified, the reliability is improved, and meanwhile, because the positions of the three small rotors are farther away from the gravity center, the provided pitching and rolling moments are larger than the mode of controlling the main rotor to tilt 13, the controllability is higher, and the wind resistance is better.

In another embodiment, front rotor 2 changes the direction of main rotor 13 tension by providing a pitching moment during mode switching for improved mode transition efficiency. The method comprises the following specific steps: when the helicopter mode is to the transition of fixed wing mode, the aircraft at first hangs down the head and leans forward, utilizes 13 pulling forces of main rotor to accelerate, and both sides wing 7 verts forward fast simultaneously and switches to the fixed wing mode, reaches main rotor 13 stall behind the settlement airspeed. When the fixed wing mode is transited to the helicopter mode, the aircraft firstly raises the head and starts the main rotor 13 to decelerate, the wings 7 on the two sides quickly tilt backwards to be switched to the helicopter mode, and the front flying speed is lowered to 0 and then the aircraft enters a hovering state.

In another embodiment, canard wings 3 are provided on both sides of the head of the fuselage frame 5 for increasing lift and moving the full-aircraft aerodynamic focus forward when operating in the fixed wing mode. A pair of canard 3 is used for providing certain lift when the flat flight and then with the effect that fixed wing mode aerodynamic focus pulled forward, improved the static stability of every single move when unmanned aerial vehicle flat flight.

In another embodiment, a horizontal tail 11 and a vertical tail 10 are provided at the tail of the fuselage frame 5, and the two ends of the horizontal tail 11 are respectively connected with the vertical tail 10 correspondingly to form an H-shaped layout. The vertical tail 10 is not provided with a rudder, only provides a vertical stabilizer to increase the course static stability in fixed-wing mode flight, and the H-shaped layout formed by the horizontal tail 11 mainly provides a horizontal stabilizer to increase the longitudinal static stability in fixed-wing mode flight.

In another embodiment, the front and rear sides of the bottom surface of the body frame 5 are provided with a nose landing gear 4 and a rear landing gear 9, respectively. The front and rear undercarriages are four-point type, have a buffering and damping structure and a large supporting area, and can ensure that the unmanned aerial vehicle has enough stability when vertically descending to touch the ground.

In another embodiment, airspeed head 1 is mounted on top of fuselage frame 5 for measuring the airspeed of a fixed-wing mode of flight. The measured airspeed is fed back to the flight control system to obtain flight state information, so that the flight is controlled accurately.

As shown in fig. 3, fig. 4 (a) - (c) and fig. 5 (a) - (c), the flight principle of the present invention is explained in conjunction with a typical mission profile:

the unmanned aerial vehicle has two flight modes of a helicopter and a fixed wing, wings tilt upwards in the take-off stage, two side rotors, the front rotor 2 and the main rotor 13 generate upward lift together, the unmanned aerial vehicle vertically takes off in the helicopter mode, after reaching a specified height, the aircraft can tilt forwards and accelerate forwards by adjusting the posture, the unmanned aerial vehicle still maintains the control mode of the helicopter mode at the moment, after reaching the transition speed, the wings 7 on two sides start to tilt forwards quickly, meanwhile, the main rotor 13 also starts to reduce the rotating speed, the unmanned aerial vehicle transits to the fixed wing mode in a short time and then slightly accelerates to reach the cruise level flight speed, the main rotor 13 completely stops rotating and is automatically locked at a feathering position parallel to the fuselage, and the whole aircraft cruises in the low-resistance level flight in the fixed wing mode. When unmanned aerial vehicle need carry out the task of hovering or need descend perpendicularly, both sides wing 7 can vert upwards fast, and the organism gesture adjustment is the hypsokinesis simultaneously, and main rotor 13 starts and rotates with higher speed, provides the reverse pulling force of deceleration and also produces partly lift simultaneously. Along with the quick tilting of the wings to the vertical state, the main rotor wing 13 and the three small rotor wings start to bear the lift force of the whole unmanned aerial vehicle, the unmanned aerial vehicle completely transits to a helicopter mode to hover and fly, and then vertical landing or transition flat flying again can be selected to be a fixed wing mode according to task requirements.

Referring to fig. 6, for different flight modes, the control method of the present invention is as follows:

helicopter mode: the flight control algorithm of the helicopter mode adopts PID control similar to a plurality of rotors; acquiring the operating state of the helicopter in a mode, such as position information and attitude information, through a navigator and a sensor; the method comprises the steps of receiving a control instruction of the ground to control the position, calculating an expected position and an expected attitude through a rigid body kinematics model to obtain an expected speed and an expected angular speed, calculating the expected speed and the angular speed through a rigid body dynamics model to obtain an expected tension and moment, generating an expected rotor wing rotating speed through a control distribution model, and finally generating an accelerator instruction through a power unit model to control.

Fixed-wing mode shape: and performing attitude control and position control on the vehicle through an L1 guidance law by adopting total energy control. Transition mode: and adopting a model-free control algorithm based on reinforcement learning to solve the problem of nonlinear aerodynamic force.

The operation modes in different states of the invention are as follows:

1. the fixed wing is in a flat flying state:

(1) Plane flight pitching operation: the level fly pitch maneuver is mainly achieved by increasing or decreasing the drag of the front propeller. And increasing the rotating speed of the propeller of the front rotor 2 to increase the tension T3 generated by the propeller, and raising the head of the unmanned aerial vehicle, otherwise lowering the head.

(2) Level flight rolling control: the level flight rolling control is mainly realized by changing the tilting angles of the left wing 7 and the right wing 7 so as to generate non-opposite lifting force. If the left wing tilts upwards to increase the attack angle, the lift force L1 of the left wing is increased, the right wing tilts downwards to reduce the lift force L2 of the right wing, and the airplane rolls rightwards.

(3) Flat flight yaw manipulation: the flat-flying yaw manipulation is mainly realized by changing the tension difference of the propulsion propellers on two sides. If increase right side rotor 12 rotational speed for right side rotor 12 pulling force T1 increases, reduces the left side rotational speed, makes 8 pulling forces T2 of left side rotor reduce, then unmanned aerial vehicle drifts about left, otherwise then drifts about right.

2. The hovering state of the helicopter is as follows:

(1) Hover pitch maneuver: the hover pitch maneuver is accomplished primarily by increasing or decreasing the differential drag of the three propellers in front of the three rotors. If the rotating speed of propellers at two ends of the wing is reduced to reduce T1 and T2, and the rotating speed of the front rotor 2 is increased to increase T3, the airplane heads up.

(2) Hovering and rolling manipulation: mainly by changing the tension difference of the small rotors at the two wingtips. If the rotor speed of the right wing tip is increased, T1 is increased, the rotor speed of the left wing tip is reduced, T2 is reduced, and the airplane can roll leftwards.

(3) The hovering yaw control is mainly realized by changing the rotating speed of three small rotors and the main rotor 13, and a part of yaw moment can also be generated by the asymmetrical tilting of wings. If the rotating speed of the main rotor 13 is increased and the rotating speed of the three small rotors is reduced, the airplane can yaw anticlockwise under the condition that the total lift force is kept unchanged. Or the right wing tilts backwards, the left wing tilts forwards, and the plane can yaw clockwise by utilizing the projection of T1 and T2 in the horizontal plane.

When the helicopter is hovered, the wings incline upwards, the reactive torque generated by the three small rotors can offset the reactive torque generated by a part of the large rotors, and meanwhile, the three small rotors can also generate upward lift force, so that the power waste caused by the installation of tail rotors is avoided. And the tail rotor can only operate the yaw of the helicopter mode, three small rotors in the invention can provide three-degree-of-freedom control moments of all flight modes, so that a rotor pitch-changing mechanism, a fixed wing pneumatic control surface, a steering engine and a large number of parts are omitted in phase change, and much weight is saved. After the aircraft is rotated into flat flight, the three small rotors can still provide the maneuvering torque and the propelling force of the aircraft body, a control surface is not needed, and the redundant weight of a maneuvering system is reduced. After the rotor completely transits to the forward flight state, the large rotor stops rotating and automatically feathers, all lift force is provided by the wings, and the lift force loss caused by the shock wave of the forward-moving blades and the stall of the backward-moving blades when the rotor flies forward is avoided.

The invention can accelerate and decelerate the forward flight in a pure helicopter mode, can perform wing tilting transition after reaching or approaching the entry speed of a fixed wing mode or can directly start the main rotor 13 to enter the helicopter mode in the fixed wing mode, can greatly shorten the transition time compared with the common tilting rotor and tilting wing aircraft, and avoids the problems that the nonlinear aerodynamic force of the transition state is difficult to predict and the control system is difficult to stabilize.

Compared with the traditional single-rotor helicopter, the invention does not need a tail rotor to consume power, does not need to design a slender tail beam, reduces the weight of a transmission system, can design a fuselage to be more compact, does not reduce a yaw control moment much, can still provide a huge yaw moment through tilting wings, and can provide a part by adjusting the rotating speed difference between the three rotors and the main rotor 13, so that the helicopter is used for maintaining balanced control under the condition of low wind speed. Compared with a double-rotor-wing tilting type airplane (such as a osprey V22), the invention uses high-speed small propellers when flying forwards, and the area ratio of the reverse flow area of the inner ring is smaller than that of the slow large propellers of the tilting rotor wing, so that the forward flying efficiency is higher. In addition, the invention can rely on the main rotor 13 to provide a part of lift force in the transition state, and the tilting process can be quickly realized, thereby avoiding that the aircraft is in a nonlinear control area for a long time. Compared with the X3 type high-speed helicopter of European helicopter company, the wing of the invention can be tilted, and the downwash of the main rotor 13 can not be shielded in a hovering state, so that the loss of lift force is small. Compared with a coaxial helicopter, the main rotor 13 of the configuration provided by the invention does not need any tilting disk and a variable pitch mechanism, the attitude control in a hovering state is completed by three small rotors, the complexity of the mechanism is simplified, the reliability is improved, and simultaneously, because the positions of the three small rotors are farther away from the gravity center, the provided pitching and rolling moments are also larger than the mode of controlling the main rotor 13 to tilt, the controllability is higher, and the wind resistance is better. In addition, when flying forward at high speed, the coaxial dual rotors generate resistance which is even more than 50% of the resistance of the whole aircraft because the hub is very high (the high design is necessary to prevent the oar from beating), but the invention does not need to tilt the main rotor 13, the main rotor 13 can be made to be closer to the aircraft body, the hub can be designed to be very short, and the resistance of the forward flying is naturally reduced greatly. Compared with a multi-rotor-fixed wing combined type unmanned aerial vehicle, the unmanned aerial vehicle has higher efficiency in hovering, and due to the fact that the huge main rotor 13 can provide large pulling force in the limit state, the effective load of the unmanned aerial vehicle can reach the same weight as the weight of the vehicle body, and the unmanned aerial vehicle can adapt to the transportation of heavy-load goods. And the aircraft can feather the main rotor 13 after turning into the flat flight state so as to achieve the purpose of reducing resistance, and the efficiency of the forward flight is not inferior to that of the multi-rotor-fixed wing combined type unmanned aerial vehicle.

The embodiments in the present description are described in a progressive manner, each embodiment focuses on differences from other embodiments, and the same and similar parts among the embodiments are referred to each other. The device disclosed by the embodiment corresponds to the method disclosed by the embodiment, so that the description is simple, and the relevant points can be referred to the method part for description.

The previous description of the disclosed embodiments is provided to enable any person skilled in the art to make or use the present invention. Various modifications to these embodiments will be readily apparent to those skilled in the art, and the generic principles defined herein may be applied to other embodiments without departing from the spirit or scope of the invention. Thus, the present invention is not intended to be limited to the embodiments shown herein but is to be accorded the widest scope consistent with the principles and novel features disclosed herein.

Claims

1. A multi-rotor combined helicopter is characterized by comprising a fuselage frame, a main rotor, wings, side rotors, a front rotor and a wing tilting mechanism;

the main rotor wing is arranged right above the fuselage frame;

the wing is connected to the fuselage frame through the wing tilting structure;

the front rotor is mounted to the head of the fuselage frame for providing a pitching maneuvering torque and providing upward lift in helicopter mode or creating a maneuvering torque by varying the difference in rotational speed with the main rotor.

2. The rotary wing composite helicopter of claim 1, wherein said wing tilt mechanism comprises a steering engine, a rudder disk sleeve and a fixed base; a clamping groove is formed in the fixed base, the steering engine is fixed in the clamping groove, and output ends on two sides of the steering engine are fixedly connected with the inner wall of the corresponding steering wheel sleeve respectively;

the fixed base is fixed inside the machine body frame;

3. The rotary wing helicopter of claim 1, wherein said rotor changes said main rotor pull direction by providing a pitch steering torque during mode switching for increasing mode transition speed.

4. The rotary wing helicopter of claim 1, wherein the side rotors are further configured to provide a roll handling torque by asymmetric speed adjustment, a yaw handling torque in cooperation with asymmetric tilting of the wing, and a yaw handling torque by increasing or decreasing the difference in rotational speed between itself and the main rotor.

5. The rotary wing helicopter of claim 1, wherein said main rotor is further configured to stall and automatically lock in a direction parallel to the fuselage upon switching to fixed wing mode.

6. The multi-rotor compound helicopter of claim 1, wherein the front rotors and the side rotors on both sides of the fuselage frame are distributed symmetrically in a triangle centered around the center of mass of the fuselage frame.

7. The multi-rotor compound helicopter of claim 1, wherein the rotation direction of said front rotor and said side rotors is opposite to the rotation direction of said main rotor for providing lift and counteracting reactive torque of said main rotor during helicopter mode.

8. The multi-rotor compound helicopter of claim 1, wherein canard wings are provided on each side of the fuselage frame head for increasing lift and moving the full aerodynamic focus forward when operating in the fixed wing mode.

9. The multi-rotor composite helicopter of claim 1, wherein the tail of the fuselage frame is provided with a horizontal tail and a vertical tail, and the two ends of the horizontal tail are respectively connected with the vertical tails to form an H-shaped layout.

10. The multi-rotor compound helicopter of claim 1, wherein a front landing gear and a rear landing gear are respectively disposed on the front and rear sides of the bottom surface of the fuselage frame. Airspeed tube is installed on fuselage frame top for airspeed when measuring machine stationary vane mode flight.