CN108528692A - Folding-wing dual-rotor aircraft and control method thereof - Google Patents

Abstract

The invention discloses a kind of folded wing double-rotor aerobat and its control methods, it is related to aircraft field, the aircraft includes fuselage, is located at the wing of fuselage both sides, two wings are hinged on fuselage and can be rotated around fuselage longitudinal axis, two ends of the wing far from fuselage are both provided with rotor system, rotor system changes rotor aerodynamic and aerodynamic torque using the feathering that built-in auto-bank unit generates, and realizes the control to the pitching of aircraft, yaw and rolling movement.The aircraft shares a set of steerable system when helicopter mode and airplane-mode work.Two sets of steerable systems are used compared to existing vertically taking off and landing flyer, the design of the present invention is easier to be reliable, aircraft no matter VTOL or it is flat fly during all without extra operating mechanism, improve the utilization rate of aircraft structural component, reduce extra construction weight.The present invention is suitable for taking photo by plane, it can also be used to which individual soldier investigates.

Description

Folding-wing dual-rotor aircraft and control method thereof

technical field

The invention relates to the field of aircraft, in particular to a double-rotor aircraft with folding wings and a control method thereof.

Background technique

Although the aerodynamic efficiency of traditional fixed-wing aircraft is relatively high, both hand-throwing and ejection have certain requirements on the venue, and hand-throwing launch requires the operator to be exposed to an open field, which is not conducive to the safety and concealment of the operator. Although quadrotor and multi-rotor aircraft are easy to take off, land and control, their flight speed, endurance time and flight radius are very limited due to the excessive flight resistance of the body, making it difficult to complete a wide range of reconnaissance tasks. The existing vertical take-off and landing aircraft utilizes the long endurance characteristics of fixed-wing aircraft and the better stability of quadrotors, and simply superimposes the functions of fixed-wing aircraft and quadrotor aircraft to achieve the purpose of vertical take-off and landing. Its working efficiency is not optimal under different flight modes, and it is difficult to give full play to the endurance potential of the aircraft. Generally speaking, the larger the diameter of the rotor, the higher the hovering efficiency. For an aircraft with the same take-off weight, the hovering efficiency of dual-rotors is higher than that of quad-rotors or multi-rotors.

To sum up, the existing vertical take-off and landing aircraft has the problems of complex structure and low aerodynamic efficiency.

Contents of the invention

Purpose of the invention: Aiming at the deficiencies and defects of the prior art, the present invention proposes a dual-rotor aircraft with folding wings, which realizes the stable manipulation of the aircraft in helicopter mode and aircraft mode through the dual-rotor system with automatic tilter. The flight does not require conventional aerodynamic rudder surfaces, and the control system in helicopter mode is still used for flight control.

Another object of the present invention is to provide a control method for the above-mentioned folding-wing dual-rotor aircraft.

Technical solution: In order to achieve the above purpose, a double-rotor aircraft with folding wings proposed by the present invention includes a fuselage and wings respectively located on both sides of the fuselage. The two wings are hinged on the fuselage and can be longitudinally around the fuselage The axis rotates, and the ends of the two wings away from the fuselage are provided with a rotor system for providing power. The rotor system uses the periodic pitch generated by the movement of the built-in automatic tilter to change the rotor aerodynamic force and aerodynamic moment, and realize the control of the pitch, yaw and roll motion of the aircraft.

The two wings are variable dihedral wings. When the aircraft is in level flight, the wings rotate around the longitudinal axis of the fuselage to change the dihedral angle, lower the center of gravity of the aircraft, and improve the stability of the aircraft. After the aircraft lands, the wings can be folded in half around the fuselage, reducing the spanwise size of the aircraft by half, making it easy to store and carry.

The rotor system includes a propeller hub, blades installed on the propeller hub, and a nacelle connected to the propeller hub. In the nacelle, a pitch-variable pull rod, an automatic tilter, a steering gear and a steering gear pull rod are arranged. The pitch-variable pull rod and the paddle The blades of the blades are connected correspondingly, the automatic tilter is connected to the variable pitch rod, the steering gear is connected to the automatic tilter through the steering gear rod, and the steering gear drives the steering gear rod to move up and down to drive the automatic tilter to move up and down and tilt, thereby driving the blade to change the pitch. distance. The automatic tilter moves up and down along the rotor shaft to change the rotor collective pitch, and the automatic tilter tilts around the spherical hinge to make the rotor produce periodic pitch change. A motor is also arranged in the nacelle, and the motor is connected with the rotor shaft to provide power to the propeller hub. The blades in the two rotor systems rotate in opposite directions, balancing the reaction torque generated during rotation.

In order to increase the directional stability, a ventral fin is provided at the tail of the fuselage, and a foldable landing gear structure is provided at the tail of the nacelle. When taking off and landing vertically, the two landing gears and the end of the ventral fin form a three-point support.

The lifting motion of the above-mentioned folding-wing dual-rotor aircraft is controlled by the rotor system. When the pulling force generated by the rotation of the rotors is greater than the total weight of the aircraft, the aircraft rises, and when it is smaller than the total weight of the aircraft, the aircraft descends. There are two ways to change the pull force of the rotor of the aircraft: one is to control the actuation direction of the steering gear to make the automatic tilter move up and down along the rotor axis when the rotor speed is constant, which drives the blade pitch to increase or decrease, thereby changing the The pulling force of the rotor; the second is to change the rotating speed of the rotor by controlling the change of the motor speed, thereby changing the pulling force of the rotor.

The tilting of the automatic tilter will cause periodic changes in the angle of attack of the blades, that is, a periodic pitch change of the rotor. When the aircraft is in the helicopter flight mode, the aircraft moves longitudinally when the longitudinal pitch is changed periodically; When the aircraft is in the aircraft flight mode, the vertical cyclical pitch control is used to control the pitch of the aircraft, the lateral cyclical pitch is used to control the heading of the aircraft, and the longitudinal cyclical pitch variable differential is used to control the roll of the aircraft.

Beneficial effects: compared with the prior art, the present invention has the following beneficial effects:

1. The double-rotor aircraft with folding wings of the present invention adopts a vertical flying wing configuration, and two rotors with automatic tilters are used as the power and control surfaces of the helicopter mode and the aircraft mode of the aircraft, which can ensure the maneuverability under the helicopter mode and stability, and can also control the aircraft in airplane mode without additional aerodynamic rudder surfaces, which simplifies the structure of the vertical take-off and landing aircraft, reduces the weight of the aircraft itself, and increases the payload. And it can actively increase stability with the help of flight control, instead of conventional aircraft mode, rely on the aircraft's own aerodynamic design to achieve static stability.

2. The aircraft of the present invention utilizes dual rotors to realize hovering/vertical take-off. After the aircraft is transferred to the aircraft mode, the lift is mainly provided by the wings, and the rotor works in the axial flow mode, which can increase the collective pitch of the rotors and reduce the speed of rotation, while ensuring that the rotors can Under the premise of providing sufficient pulling force, the drag of the rotor can be effectively reduced, and the power required by the aircraft can be reduced, so as to achieve a longer endurance time.

3. The aircraft of the present invention adopts a variable dihedral wing, which effectively lowers the center of gravity and makes the aerodynamic force on the wing more conducive to the stability of the flight of the aircraft. The variable dihedral structure operates in the helicopter mode of the aircraft, so that The center of gravity of the fuselage shifts, and the periodic pitch change of the rotor can complete the transition to the aircraft mode more efficiently, reducing the high dependence of the aircraft's tilt transition on the control efficiency of the rotor.

4. The present invention adopts a hinge folding mechanism, which can fold the aircraft in half so that the geometrical dimensions of the aircraft are greatly reduced, which is convenient for individual soldiers to carry or travel with backpacks. At the same time, the fusion design of the variable dihedral structure and the folding hinge of the aircraft not only reduces the structural complexity of the aircraft, but also reduces the structural weight of the aircraft. The variable dihedral angle feature enables the aircraft to improve flight stability by increasing the dihedral angle under unfavorable flight environment conditions, and reduces the sensitivity to the weather environment. In the helicopter mode, the center of gravity of the aircraft can be shifted by using the dihedral-changing mechanism, and the longitudinal aerodynamic force generated by the periodic variable-pitch control of the rotor can better complete the mutual conversion between the helicopter mode and the aircraft mode.

5. The invention adopts the ventral fin design, which helps to maintain the heading stability of the aircraft during the mutual switching between vertical flight and horizontal flight modes. The pelvic fins play a role in increasing the directional stability and maintaining the balance when the aircraft skids. The layout of the pelvic fin cleverly utilizes its own structure and the retractable landing gear of the nacelle to form a three-point support, which reduces unnecessary structural weight.

Description of drawings

In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the drawings that need to be used in the embodiments will be briefly introduced below. Obviously, the drawings in the following description are only some embodiments of the present invention. For Those of ordinary skill in the art can also obtain other drawings based on these drawings without making creative efforts.

Fig. 1 is the structural representation (vertical take-off) of a kind of folding wing double-rotor aircraft that the embodiment of the present invention provides;

Fig. 2 is a side view of the main structure (helicopter state) of a folding-wing dual-rotor aircraft provided by an embodiment of the present invention;

Fig. 3 is the structural representation (level flight) of a kind of folded-wing birotor aircraft provided in the embodiment of the present invention;

Fig. 4 is a schematic structural diagram of a nacelle system of a folding-wing dual-rotor aircraft provided by an embodiment of the present invention;

Figure 5 is a helicopter mode flight control strategy for a folding-wing dual-rotor aircraft provided in an embodiment of the present invention;

FIG. 6 is a flight control strategy of a folding-wing dual-rotor aircraft aircraft mode provided in an embodiment of the present invention;

Fig. 7 is a schematic diagram of the actuation of the steering gear of a folding-wing dual-rotor aircraft provided in an embodiment of the present invention.

Detailed ways

The present invention will be further described in detail below in conjunction with the accompanying drawings and embodiments. For ease of description, the drawings only show some but not all structures related to the present invention. In addition, it needs to be clear that the orientation or positional relationship indicated by "upper", "lower", "inside", and "tail" used in the terms is based on the orientation or positional relationship shown in Figure 1, and is only for description The present invention and simplified description do not indicate or imply that the referred device must have a specific orientation, be constructed and operate in a specific orientation, and thus should not be construed as limiting the present invention.

As shown in Fig. 1, the double-rotor aircraft with folding wings proposed by the present invention includes a fuselage 1 and wings 2 respectively located on both sides of the fuselage 1, and the two wings 2 are connected to the fuselage 1 by hinges and can be wound The body rotates, and the ends of the two wings 2 away from the fuselage are connected to the nacelle 6, and the head of the nacelle 6 is provided with blades 4 and fairings 5. Referring to FIG. 2 and FIG. 3 , the nacelle 6 is connected to the hub 3 , the blade 4 is installed on the hub 3 , and the blade 4 rotates with the rotation of the hub 3 . The tail of the fuselage 1 is provided with a pelvic fin 9, and the tail of the nacelle 6 is provided with two landing gear arms 7, the landing gear 8 is installed on the landing gear arms 7, and the two landing gear arms can rotate around the fulcrum inside the nacelle , implicating the retraction of the landing gear. During vertical take-off and landing, the two landing gears and the ends of the pelvic fins form a three-point support. The optoelectronic device 10 is arranged on the head of the fuselage 1 for aerial photography or investigation.

The two wings 2 are variable dihedral wings. When the aircraft is in level flight, the wings rotate around the longitudinal axis of the fuselage to change the dihedral angle, reduce the center of gravity of the aircraft, and improve the stability of the aircraft. After the aircraft lands, the wings can be folded in half around the fuselage, reducing the spanwise size of the aircraft by half, making it easy to store and carry.

Fig. 4 has shown the internal structure of nacelle 6, and nacelle 6 comprises pitch-variable pull rod 11, automatic tilter 12, steering gear 13 and steering gear pull-rod 14, and the motor 15 that is connected with rotor shaft, and pitch-variable pull rod 11 is connected with paddle The blades of blade 4 are connected in one-to-one correspondence, the automatic tilter 12 is connected with the variable distance pull rod 11, the steering gear 13 is connected with the automatic tilter 12 through the steering gear pull rod 14, and the motor 15 is used to input power to the propeller hub 3 to drive the propeller hub to rotate. The steering gear 13 drives the steering gear pull rod 14 to actuate, so that the automatic tilter 12 moves up and down or tilts, and drives the blades 4 to change the pitch. When all the steering gears in the nacelle 6 have the same momentum, the automatic tilter 12 moves up and down along the rotor shaft to drive the angle of attack of the blades 4 to increase or decrease, thereby changing the collective pitch of the rotor. When the amount of momentum of the steering gear is different, the automatic tilter 12 will tilt around the spherical hinge, so that the angle of attack of the blade 4 will change periodically, that is, the periodic pitch of the rotor will be changed, so that the blade 4 will produce the angle of attack at different rotation azimuth angles. The lift is different. The blades are generally made of elastic materials. When the lift force received by different azimuths is different, the deformation of the blades in the waving direction is also different, that is, the rotor paddle will tilt, and the pulling force generated by the rotor will have a direction in which the paddle is tilted. portion.

The lifting motion of the aircraft is controlled by the rotor system. When the pulling force generated by the rotation of the rotor is greater than the total weight of the aircraft, the aircraft rises, and when it is less than the total weight of the aircraft, the aircraft descends. There are two ways to change the pull force of the rotor of the aircraft: one is to control the actuation direction of the steering gear to make the automatic tilter move up and down along the rotor axis when the rotor speed is constant, which drives the blade pitch to increase or decrease, thereby changing the The pulling force of the rotor; the second is to change the rotating speed of the rotor by controlling the change of the motor speed, thereby changing the pulling force of the rotor.

When the aircraft takes off vertically, the aircraft is lifted to a certain height by the pulling force generated by the dual rotors (i.e., the two blades 4), and the two landing gears 8 are respectively retracted into the interior of the corresponding nacelle 6, and the automatic tilter 12 is used to control the movement of the aircraft in all directions. Movement and stability, reaching a safe height. As mentioned above, the periodic pitch change of the automatic tilter causes the rotor disc to generate a lateral force, forming a tilting moment on the fuselage, and the two wings 2 rotate around the fuselage 1, so that the center of gravity of the aircraft is shifted to the tilting side, and the rotor The inclination of the propeller disk makes the aircraft accelerate in the horizontal direction. With the increase of the forward flight speed, the lift generated by the two wings gradually increases, and the power consumption of the rotor gradually decreases until the lift of the aircraft is completely provided by the two wings. To generate the pulling force of the forward flight and control the flight attitude of the aircraft to stabilize, it is possible to increase the collective pitch of the rotors and reduce the rotational speed of the rotors, so that the required power of the rotors can be reduced under the condition of constant pulling force, and the endurance time of the aircraft can be increased. When the aircraft needs to land, it needs to use the automatic tilter of the rotor to cooperate with the rotation of the two wings to change the aircraft from a horizontal flight attitude to a vertical flight attitude. Torque, when the angle of attack of the aircraft increases, the resistance of the aircraft will also increase, that is, the aerodynamic braking effect will help the aircraft to decelerate. At this time, it will switch to the helicopter mode, and the aircraft will be controlled to land accurately according to the set landing point to complete the flight mission. .

When the aircraft is in the helicopter mode, the aircraft will move vertically when the longitudinal pitch is changed periodically; when the horizontal pitch is changed, the aircraft will move laterally; when the longitudinal pitch is changed differentially, the aircraft will change its heading. In the aircraft mode, the vertical cyclical pitch control is used to control the pitch of the aircraft, the lateral cyclical pitch is used to control the heading of the aircraft, and the longitudinal cyclical pitch variable differential is used to control the roll of the aircraft.

As shown in Figure 5, in the helicopter mode, it is possible to control the total rotor pitch, motor speed, longitudinal cyclic pitch, longitudinal cyclic pitch differential, and lateral cyclic pitch. The collective pitch and motor speed need to be controlled independently. As shown in Figure 6, in the aircraft mode, it is possible to directly control the collective pitch, motor speed, longitudinal cyclical pitch change, longitudinal cyclical pitch variable differential, and lateral cyclical pitch change. The control of the collective pitch of the rotor is realized by moving the automatic tilter up and down along the rotor shaft to increase or decrease the pitch of the blades, and the control of the motor speed is realized by the adjustable pulse width signal sent by the flight control system. Variations can change the rotor pull, which controls the lift motion of the aircraft. The control of vertical and horizontal periodic pitch is realized by controlling the movement of the automatic tilter, and the movement of the automatic tilter is realized by controlling the actuating direction of the steering gear. Figure 7(a)-7(h) shows the action of the steering gear Schematic diagram of the operation, the control number in the figure indicates the different driving positions of the steering gear, among which, the steering gear 1-3 is located in one nacelle, and the steering gear 4-6 is located in the other nacelle, and the direction of the arrow indicates the driving direction of the steering gear. Figure 7(a)-7(h) respectively show that different motions of the aircraft correspond to different automatic tilter inputs, Figure 7(a) is a schematic diagram of raising the collective distance, Figure 7(b) is a schematic diagram of reducing the collective distance, Figure 7 (c) Schematic diagram of longitudinal periodic pitch change (rear tilt), Figure 7(d) schematic diagram of longitudinal periodic pitch change (forward tilt), Figure 7(e) schematic diagram of longitudinal periodic pitch variable differential (left turn), Figure 7(f) Schematic diagram of longitudinal periodic pitch variable differential (right turn), Figure 7(g) schematic diagram of lateral periodic pitch variable (right bias), and Figure 7(h) schematic diagram of lateral periodic pitch variable (left bias). Take Figure 7(c) as an example, to make the aircraft tilt backward, the No. 1 and No. 4 steering gears are driven upwards, and No. 2, 3, 5, and 6 steering gears are driven downwards. Similarly, in other cases, the driving direction of the steering gear can be obtained by referring to the corresponding diagram, and details will not be repeated one by one.

Claims (9)

1. A double-rotor aircraft with folding wings, is characterized in that, comprises fuselage, the wing that is respectively positioned at described fuselage both sides, and described two wings are hinged on fuselage and can rotate around fuselage longitudinal axis , the ends of the two wings far away from the fuselage are provided with a rotor system for providing power. 2. A kind of folding-wing dual-rotor aircraft according to claim 1, wherein the rotor system comprises a hub, blades mounted on the hub, and a nacelle connected to the hub , the nacelle is provided with a pitch-variable pull rod, an automatic tilter, a steering gear and a steering gear pull-rod, the pitch-variable pull-rod is correspondingly connected to the blade of the blade, and the automatic tilter is connected to the pitch-variable pull rod, The steering gear is connected to the automatic tilter through a steering gear rod, and the steering gear drives the steering gear rod to move up and down to drive the automatic tilter to move up and down and tilt, thereby driving the blades to change pitch. 3. A kind of folding-wing dual-rotor aircraft according to claim 2, characterized in that, the nacelle is also provided with a motor for providing power to the hub, and the motor is connected with the rotor shaft. 4. A kind of folding wing dual-rotor aircraft according to claim 2, is characterized in that, the rotation directions of the blades in the two rotor systems are opposite. 5. A kind of folding-wing dual-rotor aircraft according to claim 2, characterized in that, a ventral fin is provided at the tail of the fuselage, and a foldable landing gear structure is provided at the tail of the nacelle. , the two landing gears and the ends of the pelvic fins form a three-point support. 6. The double-rotor aircraft with folding wings according to claim 1, wherein the two wings are variable dihedral wings. 7. A control method for the folding-wing dual-rotor aircraft according to any one of claims 1-6, wherein the lifting motion of the aircraft is controlled by the rotor system, and the pulling force generated by the rotation of the rotor is greater than that of the aircraft The aircraft rises when the total weight is less than the total weight of the aircraft, and the aircraft descends; the flight motion of the aircraft in the helicopter mode and the airplane mode is controlled by the rotor cycle pitch change. 8. The control method of the folding-wing dual-rotor aircraft according to claim 7, wherein the method for changing the pulling force of the aircraft rotor comprises: When the rotor speed is constant, the automatic tilter moves up and down along the rotor shaft by controlling the actuating direction of the steering gear, which drives the blade pitch to increase or decrease, thereby changing the rotor tension; and By controlling the change of the rotational speed of the motor to drive the rotor to change the rotational speed, thereby changing the pulling force of the rotor. 9. The control method of the folding-wing dual-rotor aircraft according to claim 7, characterized in that, in the helicopter mode, when the vertical cycle pitch is changed, the aircraft moves longitudinally; when the horizontal cycle pitch is changed, the aircraft moves laterally; Periodic variable pitch differential, the aircraft changes its heading; In the aircraft mode, the vertical cyclical pitch control is used to control the pitch of the aircraft, the lateral cyclical pitch is used to control the heading of the aircraft, and the longitudinal cyclical pitch variable differential is used to control the roll of the aircraft.