Four-flapping-wing aircraft and attitude control system thereof

Abstract

The invention discloses a four-flapping-wing aircraft and an attitude control system thereof, and relates to the technical field of micro aircrafts, and the four-flapping-wing aircraft comprises a fixed rack, a guide rod, a slide block mechanism, a pitching frame, a rolling frame, a pulley and a chord driving system. The fixed rack is provided with four cantilevers, a sliding block mechanism is installed on a guide rod, and the pitching frame and the rolling frame control pitching and rolling actions through steering engines respectively. The four pulleys are connected to the pitching frame, and each pulley is hinged to one wing. The string driving system comprises a guide wheel, a string and a crank, one end of the string is connected with the slider mechanism, and the other end bypasses the guide wheel and is connected with the pulley to realize wing flapping. Through the deflectable spars and fine steering engine control, efficient aerodynamic performance, enhanced thrust output and excellent flight flexibility and adaptability are achieved, meanwhile, due to the design of the crank sliding blocks, structural stress is relieved, the service life is prolonged, the aircraft adapts to multiple flight modes, and the environment adaptability and task execution efficiency of the aircraft are improved.

Description

Technical Field

[0001] The invention belongs to the technical field of micro-aircraft, and relates to a bionic flapping-wing aircraft, and in particular to a four-flapping-wing aircraft and a posture control system thereof. Background Art

[0002] Flapping Wing Aerial Vehicle (FWAV) is a bionic aircraft that imitates the flapping wings of flying animals in nature (such as birds, insects, etc.). Due to its light and flexible characteristics, this type of aircraft has shown great potential in indoor reconnaissance, environmental detection, scientific and educational research, etc. However, the existing flapping wing aircraft technology still has many problems in terms of drive system and attitude control, which limits its wide application and development.

[0003] In terms of the drive system, existing flapping-wing aircraft mainly use a motor-driven crank-chord mechanism, which will generate great stress when the wings are rapidly accelerated and decelerated, leading to mechanical fatigue and performance degradation. In addition, the energy efficiency of this drive method is not high, especially during the ascent phase of the wings, due to air resistance, the energy loss is large. For large or multi-wing structure flapping-wing aircraft, the complexity and uneven force problems of traditional drive systems are more obvious.

[0004] In terms of attitude control, the stable flight of existing flapping-wing aircraft requires precise attitude control, which is currently mainly achieved by changing the center of gravity, adjusting the wing root torsion angle, or using a multi-axis attitude control system. These methods have their own advantages and disadvantages, such as low center of gravity control accuracy, complex wing root torsion angle control system, and multi-axis attitude control system, although accurate, may have delayed response.

[0005] In summary, although the flapping-wing aircraft with a four-wing structure can provide greater thrust and stability, it has higher requirements for the control system and traditionally relies on complex multi-axis control systems, which increases the complexity of hardware and software and may cause attitude adjustment lag. For this reason, it is necessary to design a new drive method for a four-wing flapping-wing aircraft based on a string and crank slider mechanism, as well as an attitude control system combining a servo and a linear servo, to effectively relieve mechanical stress and improve the accuracy and response speed of attitude control. Summary of the invention

[0006] The purpose of the present invention is to provide a four-flapping-wing aircraft and its attitude control system, which solves the problems of low driving efficiency, slow attitude control response and stress concentration of the driving system existing in traditional flapping-wing aircraft by introducing a deflectable wing beam design and a simplified attitude control system and a crank slider system, thereby improving the overall flight performance, aerodynamic efficiency and service life of the aircraft.

[0007] Specifically, the present invention aims to achieve the following objectives:

[0008] 1. Improve aerodynamic performance; by designing a deflectable wing beam structure at the root of the wing, the flapping-wing aircraft can passively adjust the angle of attack with the airflow during flight, thereby improving the aerodynamic efficiency of the wing, increasing thrust output, and improving the flexibility and adaptability of the aircraft.

[0009] 2. Alleviate the stress concentration problem in the drive system and wing acceleration and deceleration process; by combining the drive system of the chord and crank slider mechanism, the stress during acceleration or deceleration is reduced, the service life is extended, and the reliability is improved.

[0010] 3. Improve the response speed and accuracy of attitude control; use a combination of servos and linear servos to quickly control pitch, roll and yaw torque by adjusting the inclination of the wing flapping plane, simplify the structure and reduce delays.

[0011] 4. Enhance the adaptability of flapping-wing aircraft to complex flight environments; involve a control system based on the change of the wing flapping plane angle, improve the response speed, effectively deal with sudden airflow and narrow space, and ensure flight stability and safety.

[0012] 5. Simplify the system structure and reduce weight; by simplifying the control system and reducing weight, design a system structure suitable for micro-aircraft to achieve lightweight and miniaturization.

[0013] 6. Achieve the coupling effect of roll control; by coupling the pitch and roll attitude of the front and rear wing blades, simplify the yaw control and achieve efficient lateral attitude adjustment, which is suitable for high-maneuverability flight.

[0014] In general, the present invention aims to design a flapping-wing aircraft that can fly efficiently in complex environments, quickly and accurately control attitude, and has a light and stable structure. Through an improved drive system and attitude control system, the shortcomings of the existing technology are addressed and a more efficient and reliable flight solution is provided.

[0015] To achieve the above object, in one aspect, the present invention provides a four-flapping-wing aircraft, comprising:

[0016] A fixed frame; a guide rod is provided in the middle of the fixed frame along the front-rear direction, a slider mechanism is reciprocatingly slid on the guide rod, and the fixed frame has four cantilevers extending in the front-rear direction;

[0017] Two pitch frames are symmetrically connected to the front and rear ends of the fixed frame, and the pitching motion of the two pitch frames is controlled by a straight-line steering gear;

[0018] A roll frame is connected to the outside of the pitch frame in a rotational direction, and the roll frame is controlled by a rotary servo to drive the front and / or rear wings to roll left and right;

[0019] There are four pulleys, which are horizontally rotatably connected to the left and right ends of the two pitch frames, and a wing is hingedly connected to the outer peripheral side of each pulley;

[0020] A string driving system, the string driving system includes guide wheels rotating on the four cantilevers, strings connecting the slider mechanism and the pulley, and a crank driving the slider mechanism to reciprocate; one end of the string is fixed to one side of the slider mechanism, and the other end is wrapped around the guide wheel and connected to the pulley; two strings are wrapped around each pulley in the forward and reverse directions, and the two strings are respectively connected to the front and rear sides of the slider mechanism through different guide wheels; the crank is connected to a power source through a transmission system.

[0021] The above structure aims to propose a four-wing aircraft, the design principle of which is to design a wing beam that allows slight deflection at the root of the wing, so that the wing can passively adjust its angle according to the change of airflow during flight. This design imitates the behavior of bird wings in nature, allowing the wings to automatically adjust when the airflow pressure or direction changes to maintain the best flight state. Because the structure of the wing root allows slight deflection, the wing of the flapping-wing aircraft will automatically adjust the angle of attack during each up and down flapping of the wing. Especially when moving quickly or facing sudden airflow, this structure can effectively optimize the relative angle between the wing and the airflow, increasing lift and thrust output.

[0022] Optionally, wing fixing arms are symmetrically extended from left and right sides of the pitch frame, and the wing fixing arms are rotatably connected to the pulley.

[0023] Optionally, sliding limiter fixing arms are symmetrically extended from the left and right sides of the rolling frame, a sliding limiter is fixedly connected to the sliding limiter fixing arms, and the sliding limiter is movably connected to the wing to change the flapping plane of the wing.

[0024] Optionally, a wing support beam and a wing root deflection beam are respectively provided on the upper edge and the side edge of the wing, and one end of the wing support beam passes through the sliding limiter and is hinged to the peripheral side of the pulley.

[0025] Optionally, the wing root deflection beam is hinged to the end of the wing fixing arm through a second hinge structure.

[0026] Optionally, the linear servo is fixedly connected below the fixed frame, and a movable block is provided on one side of the servo towards the pitch frame for reciprocating up and down, a connecting rod is hinged on the movable block, and one end of the connecting rod away from the movable block is hinged on the pitch frame.

[0027] Optionally, the rotary servo is fixed in the pitch frame, and a rudder cap is fixedly connected to the output shaft of the rotary servo. The rudder cap is an equal-arm lever structure, the center of which is fixedly connected to the output shaft of the rotary servo, and both ends are hinged to the rolling frame through a connecting shaft.

[0028] Optionally, the wings are distributed in a rectangular shape and have a flapping amplitude of 120°.

[0029] Optionally, the power source includes a motor and a power supply, and the motor and the power supply are fixed at the bottom center of the fixed frame.

[0030] On the other hand, the present invention provides a four-flapping-wing aircraft attitude control system, based on the four-flapping-wing aircraft described in any of the above items, the system includes: real-time collection of attitude data of the four-flapping-wing aircraft, and sending control signals to the motor and four servos according to preset attitude expected values; controlling the lift by the rotation speed of the motor to achieve control of the flight altitude; controlling the roll frame by the servo to change the angles of the left and right wings to control the roll of the aircraft on the horizontal plane; controlling the pitch frame by the servo to change the pitch angles of the front and rear pairs of wings to make the aircraft pitch up and down; and achieving yaw control by the difference in the lateral and longitudinal flapping planes between the four wings.

[0031] The present invention discloses the following technical effects:

[0032] The present invention proposes an efficient and flexible flapping-wing aircraft through innovative structural design and drive control scheme, which has the following significant technical effects:

[0033] 1. Improved aerodynamic performance. The deflectable wing beam design allows the wing to produce a larger angle of attack during flapping, significantly improving the aerodynamic performance of the flapping-wing aircraft. The deflected wing beam can dynamically adjust the angle of the wing relative to the airflow, enhancing the adaptability of the wing in different flight phases. Compared with the traditional fixed wing beam design, the wing beam of the present invention reduces airflow resistance, improves the thrust output of the aircraft, and achieves more efficient flight. This design improves the interaction between the flapping-wing aircraft and the airflow, especially when facing sudden airflow or environmental changes, the deflection of the wing beam can effectively reduce vibration and improve flight stability.

[0034] 2. The drive system is more durable and mechanical stress is greatly reduced. The present invention adopts a drive system that combines a chord with a crank slider mechanism, which effectively alleviates the stress concentration problem that occurs in traditional flapping-wing aircraft during wing acceleration and deceleration. Through the slider mechanism, the mechanical load caused by the rapid acceleration and deceleration of the wing can be dispersed, which significantly reduces mechanical fatigue and wear and extends the service life of the system. Compared with the traditional direct mechanical transmission method, the present invention optimizes the power transmission method through the combination of chord transmission and slider mechanism, making the operation of the entire system more stable.

[0035] 3. Fast attitude control response speed and precise control. The present invention realizes fast and precise flight attitude adjustment through an attitude control system based on plane angle changes. Compared with the traditional complex multi-axis attitude control system, the design of the present invention not only simplifies the structure of the control system, but also significantly improves the control accuracy and response speed, can quickly respond to sudden airflow or external environment changes, and provide faster attitude adjustment.

[0036] 4. Simplified roll control and flexible attitude adjustment. The present invention realizes yaw control through the difference in pitch angles between the front and rear wings and roll coupling, and no longer relies on the traditional complex multi-axis system. This design greatly simplifies the control system while maintaining flexible attitude adjustment capabilities, especially providing a smooth control experience when performing complex flight maneuvers.

[0037] 5. Enhanced flight stability and stronger ability to adapt to the environment. The combination of the drive system and attitude control system of the present invention enables the flapping-wing aircraft to respond quickly and accurately to various environmental changes during flight, especially when responding to sudden changes in airflow and environmental interference encountered during flight, and can effectively maintain flight stability. Through the mutual cooperation of the deflectable wing beam, the slider mechanism, and the attitude control system based on the change of plane angle, the flapping-wing aircraft of the present invention exhibits excellent adaptability in complex environments, reducing the situation of flight attitude loss of control due to changes in the external environment. BRIEF DESCRIPTION OF THE DRAWINGS

[0038] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0039] Figure 1 A front view of a four-flapping-wing aircraft provided by an embodiment of the present invention;

[0040] Figure 2 It is a right side view of the four-flapping-wing aircraft of the present invention;

[0041] Figure 3 A top view of the four-flapping-wing aircraft of the present invention;

[0042] Figure 4 is an axonometric view of the four-flapping-wing aircraft of the present invention;

[0043] Figure 5 The structure diagram of the frame assembly part of the four-wing aircraft of the present invention is shown in FIG. Figure 1 ;

[0044] Figure 6The structure diagram of the frame assembly part of the four-wing aircraft of the present invention is shown in FIG. Figure 2 ;

[0045] Figure 7 The bottom structure diagram of the frame assembly part of the four-flapping-wing aircraft of the present invention;

[0046] Figure 8 A diagram showing the connection relationship between the pitch frame and the roll frame in the four-flapping-wing aircraft of the present invention;

[0047] Fig. 9 for Figure 8 Exploded view of;

[0048] Fig.10 An exploded view of the frame assembly part of the four-flapping-wing aircraft of the present invention;

[0049] Fig.11 It is a structural schematic diagram of a conventional system of fixing a frame and gears in a four-flapping-wing aircraft of the present invention;

[0050] Fig.12 It is a schematic diagram of the structure of the conventional gear system in the four-flapping-wing aircraft of the present invention;

[0051] Fig.13 It is a schematic diagram of the structure of the chord drive system in the four-flapping-wing aircraft of the present invention;

[0052] Fig.14 It is a schematic diagram of the flapping state of the wings of the four-wing aircraft of the present invention;

[0053] Fig.15 It is a schematic diagram of the rolling attitude control of the four-flapping-wing aircraft of the present invention;

[0054] Fig.16 It is a schematic diagram of pitch attitude control of a four-flapping-wing aircraft of the present invention;

[0055] Fig.17 Schematic diagram of yaw attitude control of a four-flapping-wing aircraft of the present invention;

[0056] Fig.18 It is a schematic diagram of attitude control of the wing root deflection beam in the four-flapping wing aircraft of the present invention;

[0057] In the figure: 1, wing; 2, frame assembly; 3, steering gear 1; 4, motor; 5, power supply; 6, rudder cap; 7, connecting rod; 8, steering gear 3; 9, steering gear 4; 10, guide wheel; 10a, first guide wheel; 10b, second guide wheel; 11, pulley; 11a, first pulley; 11b, second pulley; 12, sliding stopper; 13, steering gear 2; 14, rolling frame; 15, guide rod; 16, slider mechanism; 17, chord; 1 7a, first chord; 17b, second chord; 17c, third chord; 17d, fourth chord; 18, wing support beam; 19, wing root deflection beam; 20, pitch frame; 21, driving gear; 22, speed gear; 23, driven gear; 24, fixed frame; 25, crank; 26, pitch axis; 27, roll axis; 28, wing fixing arm; 29, sliding limiter fixing arm; 30, hinge mechanism one; 31, hinge mechanism two. DETAILED DESCRIPTION

[0058] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0059] The chord drive system of flapping-wing aircraft in the prior art mostly adopts a crank-chord mechanism driven by a motor. The principle of this system is to drive the chord to pull the wing through the rotation of the crank, so that it flaps up and down to generate thrust. However, there are two main problems with this driving method: one is stress concentration. During the rotation of the crank, the acceleration and deceleration of the wing are extremely rapid, especially at the top and bottom of the motion cycle, the chord will generate a huge sudden tension and stress. This stress concentration is easy to cause mechanical fatigue of the transmission system, affecting the overall life and performance of the flapping-wing aircraft; the second is low energy efficiency. The traditional crank-chord drive has a relatively low energy utilization rate when the wing moves up and down. During the flight, the rising stage of the wing cannot provide effective thrust due to the large air resistance, resulting in high energy loss. In addition, for flapping-wing aircraft with larger wingspans or multi-wing structures, the complexity and uneven force of the traditional crank transmission system are more prominent. To solve this problem, there is an urgent need for a more optimized driving method that can reduce stress concentration and transmit motion smoothly and efficiently.

[0060] The attitude control of flapping-wing aircraft is the key to achieving stable flight. At present, the common attitude control of flapping-wing aircraft mainly relies on the following methods:

[0061] Center of gravity change control: By changing the center of gravity position of the aircraft, its overall flight attitude is affected. Although this method is simple, the control accuracy is low, especially in fast flight, it is difficult to make subtle adjustments.

[0062] Wing root twist angle control: Some flapping-wing aircraft use the twist angle at the wing root to affect the flapping angle of the wing, thereby adjusting the flight attitude. This method is flexible to control, but requires a more complex multi-axis motion mechanism, and the system design and debugging are difficult, which can easily lead to feedback delays and unstable control.

[0063] Multi-axis attitude control: Some flapping-wing aircraft have introduced multi-axis inertial sensors, combined with motors or servo systems to adjust the roll, pitch, and yaw angles of the aircraft. Although this method can achieve relatively precise control, the system structure is complex and easily increases the weight and power consumption of the aircraft. At the same time, more complex sensing and control systems also have a certain response delay, and the reaction speed is not ideal when dealing with sudden airflow or changes in the external environment.

[0064] Most existing quadcopter aircraft rely on complex multi-axis control systems to achieve attitude adjustment in three-dimensional space. This control system not only increases the complexity of hardware and software, but also easily causes attitude adjustment lag and is difficult to cope with sudden airflow changes. At the same time, existing technologies also have problems such as mechanical stress, control complexity, low energy efficiency, and difficulty in multi-wing coordination.

[0065] To sum up, in view of the limitations of existing flapping-wing aircraft in terms of drive systems and attitude control, the purpose of the present invention is to provide a four-flapping-wing aircraft and its attitude control system, which aims to effectively alleviate mechanical stress, improve the accuracy and response speed of attitude control, and especially show strong control advantages in the design of multi-wing structures.

[0066] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0067] Reference Figures 1 to 18 As shown, an embodiment of the present invention provides a four-flapping-wing aircraft, including a frame assembly 2 and wings 1 movably connected to the four corners of the frame assembly 2, wherein the frame assembly 2 is composed of a fixed frame 24, a roll frame 14, a pitch frame 20, a gear conventional system, a chord drive system, a power source and multiple attitude control servos.

[0068] In a specific embodiment, the fixed frame 24 provides the necessary structural support for the aircraft, ensuring that the various components are connected and assembled to each other and operate in the correct position, such as Figure 7 and Figure 8As shown, the fixed frame 24 includes four cantilevers that are symmetrical in pairs. The four cantilevers are grouped in pairs and extend in the same direction (front and rear), and the two cantilevers in the same group are integrally formed and made of carbon fiber or composite materials, with sufficient strength and rigidity to withstand the force during movement. In some embodiments, the fixed frame 24 can be cut from carbon fiber plates and 3D printed with resin materials, so that the aircraft fuselage has a lighter weight while ensuring its structural strength. The two groups of cantilevers are connected by two guide rods 15, and two slider mechanisms 16 slide on the guide rods 15. The slider mechanism 16 is used to fix the chord 17. The slider mechanism 16 performs linear reciprocating motion along the guide rod 15 under the drive of the gear conventional system, thereby driving the chord 17 to move. The chord 17 transmits the reciprocating motion to the wing 1 through the guide wheel 10 and the pulley 11 in the chord driving system, driving the wing 1 to flap the wing 1. A rotating shaft is provided at the far end of the four cantilevers of the fixed frame 24, and the rotating shaft is used to rotate and connect the guide wheel 10. The position of the guide wheel 10 and the fixed frame 24 is fixed, and the slider mechanism 16 reciprocates on the guide rod 15 of the fixed frame 24. One end of the string 17 is fixed to one side of the slider mechanism 16, and the other end is fixed to the outer peripheral side of the pulley 11 after passing the guide wheel 10. The pulley 11 is used to fix the wing support beam 18 of the wing 1. In the above structure, the slider mechanism 16 is connected to the string 17, and the driving force is converted into the pulling action of the string 17. The guide wheel 10 and the pulley 11 are used to change the transmission direction of the string 17, so that the force transmission is smoother and more accurate. The string 17 is a flexible transmission component, which transmits the mechanical force generated by the crank 25 to the wing 1, pulls the wing 1 to complete the flapping, and the wing 1 is pulled by the string 17 to swing and complete the flapping flight action.

[0069] In a specific embodiment, if Figures 5 to 9 As shown, shaft seats are arranged facing each other on the two cantilevers in each group of the fixed frame 24, and the two shaft seats are rotatably connected to the pitch frame 20 through the pitch axis 26. Roll axes 27 are arranged at the front and rear ends of the pitch frame 20. The roll frame 14 is a square frame, and the roll frame 14 is located on the outside of the pitch frame 20. The center of the front and rear side frames is rotatably connected to the roll axis 27. The left and right sides of the roll frame 14 are fixedly connected to sliding limiter fixing arms 29, and the end of the sliding limiter fixing arms 29 is fixedly connected to a sliding limiter 12 for limiting the movable position of the wing support beam 18. The sliding limiter 12 is a semicircular frame structure, which is half surrounded by the outside of the pulley 11, and a sliding limiter groove for accommodating the wing support beam 18 passing through is formed in the middle of the sliding limiter 12, as shown in FIG. Fig. 9 shown.

[0070] In a specific embodiment, if Fig. 9 As shown, the pitch frame 20 includes a horizontal square frame body, and a steering gear 3 ( Fig. 9Taking one side of the fixed servo 13 as an example, it should be understood that the pitch frame 20 on the other side of the fixed servo 2 13 has the same structure. One side of the pitch frame 20 is bent upward at 90 degrees and connected to the pitch axis 26 to achieve connection with the fixed frame 24. The left and right sides of the pitch frame 20 extend outward with wing fixing arms 28, which are located above the sliding limiter fixing arms 29, and the ends of which are used to install the pulley 11. The pulley 11 is rotatably connected to the upper end of the wing fixing arm 28. The circumferential surface of the pulley 11 is fixed with a hinge mechanism 130, and the wing support beam 18 is hingedly installed on the hinge mechanism 130, and the sliding limiter 12 is sleeved outside the wing support beam 18 to limit the wing support beam 18. The end of the wing fixing arm 28 is also fixedly connected with a hinge mechanism 2 31 , on which a wing root deflection beam 19 is hingedly mounted. The wing support beam 18 and the wing root deflection beam 19 are supported on the upper edge and the side edge of the wing 1 respectively.

[0071] In some optional embodiments, the deflectable wing support beam 18 is connected to the fuselage through a hinge mechanism 30, allowing free deflection within a limited angle. In order to ensure that the deflection angle is moderate, a damper can be set at the hinge mechanism 30 so that the swing of the wing support beam 18 has a certain damping property to control the free deflection amount of the wing 1 and avoid excessive angles affecting flight stability.

[0072] Similarly, in some optional embodiments, a damper is also provided between the wing root deflection beam 19 and the hinge mechanism 2 31, so that the deflection of the wing root deflection beam 19 has a certain damping property, such as Fig.18 As shown, the root of the wing 1 of the embodiment of the present invention adopts a deflectable wing beam design, which improves the adaptability of the wing 1.

[0073] In a specific embodiment, if Figure 8 As shown, the servo 13 is fixed in the pitch frame 20, and its output shaft faces the front end, in the same direction as the roll axis 27. The output shaft is fixedly connected to the rudder cap 6, and the middle part of the rudder cap 6 is fixedly connected to the output shaft of the servo 13. The two sides are connected to the roll frame 14 through the connecting shaft, and the positions connected to the roll frame 14 are located on the opposite sides of the roll axis 27. When the servo 13 is running, it can drive the rudder cap 6 to change its angle, and then drive the roll frame 14 to flip around the roll axis 27 to achieve the control of the rolling attitude of the aircraft. By adjusting the angle difference of the flapping planes of the wings 1 on the left and right sides, the rolling flight attitude can be changed. It should be understood that the servo 2 13 located at the rear side of the aircraft can achieve the same attitude control effect.

[0074] In a specific embodiment, Fig.13As shown, the string drive system includes a slider mechanism 16 sliding on a guide rod 15, and strings 17 fixed to both sides of the slider mechanism 16. There are four strings 17 in total. One end of the string 17 is fixed to one side of the slider mechanism 16, and the other end is wound around the guide wheel 10 and fixed on the pulley 11. The outer peripheral sides of the guide wheel 10 and the pulley 11 are both provided with two upper and lower winding grooves. Fig.13 As shown, the guide wheel 10 at the front end is defined as the first guide wheel 10a, the guide wheel 10 at the rear end is defined as the second guide wheel 10b, the pulley 11 at the front end is defined as the first pulley 11a, and the pulley 11 at the rear end is defined as the second pulley 11b; the first string 17a passes through the upper winding groove of the first pulley 11a and then is wound counterclockwise in the upper winding groove of the first pulley 11a, the second string 17b passes through the lower winding groove of the first pulley 11a and then is wound counterclockwise in the lower winding groove of the second pulley 11b, the third string 17c passes through the upper winding groove of the second pulley 11b and then is wound clockwise in the upper winding groove of the second pulley 11b, and the fourth string 17d passes through the lower winding groove of the second pulley 11b and then is wound clockwise in the lower winding groove of the first pulley 11a. In the string drive system, the first string 17a and the second string 17b are fixed to the front side of the slider mechanism 16, both of which bypass the first guide wheel 10a and are connected to the first pulley 11a and the second pulley 11b respectively; the third string 17c and the fourth string 17d are fixed to the rear side of the slider mechanism 16, both of which bypass the second guide wheel 10b and are connected to the second pulley 11b and the first pulley 11a respectively. The four strings 17 cross-connect the slider mechanism 16 and the four wings 1 to ensure that each wing 1 is subjected to balanced and symmetrical tension during reciprocating motion.

[0075] In some optional embodiments, the string 17 is made of high-strength material, which can withstand the tensile force in high-speed movement. For example, the string 17 can be made of polyurethane, polyamide, chloroprene rubber and other materials with light weight, high strength and high wear resistance.

[0076] In a specific embodiment, the attitude control servo includes servo 1 3 and servo 3 8 located at the front side of the frame assembly 2, and servo 2 13 and servo 4 9 located at the rear side of the frame assembly 2, wherein servo 1 3 and servo 2 13 are used to control the roll of the front and rear sets of wings 1. Specifically, servo 1 3 and servo 2 13 are responsible for controlling the roll attitude of the aircraft. The servo changes the angle of the wings 1 on the left and right sides by controlling the roll frame 14 to control the roll of the aircraft on the horizontal plane. Servo 3 8 and servo 4 9 are used to control the pitch of the front and rear sets of wings 1. Specifically, servo 3 8 and servo 4 9 are used to control the pitch attitude of the aircraft. By adjusting the pitch angle of the front and rear pairs of wings 1, the aircraft can perform up and down pitch movements.

[0077] The connection method of the servo 1 3 and the servo 2 13 to the rolling frame 14 is detailed in the above description of the structure of the rolling frame 14 and the pitching frame 20, which will not be repeated here.

[0078] The connection relationship between the servo 3 8 and the servo 4 9 and the pitch frame 20 is as follows: Figure 6 As shown, the servo three 8 and the servo four 9 are both fixedly installed below the fixed frame 24. Taking the servo three 8 as an example, the servo is a straight-acting servo, and a movable block is provided on the side facing the pitch frame 20 to reciprocate up and down. The movable block is hinged to the connecting rod 7, one end of the connecting rod 7 is hinged to the movable block, and the other end is hinged to the pitch frame 20. When the servo three 8 is running, it drives the movable block to move up or down, and then drives the pitch frame 20 to change the pitch angle through the connecting rod 7, thereby realizing the control of the flight attitude of the wing 1.

[0079] In a specific embodiment, the gear transmission system and the power source are placed in the central area of ​​the fuselage, that is, at the bottom of the fixed frame 24, and the string drive system is installed on both sides of the fixed frame 24 to form a balanced layout to ensure symmetry and stability during movement. The gear transmission system includes a driving gear 21, a speed gear 22 and a driven gear 23, wherein the driving gear 21 is fixed to the output shaft of the motor 4, the large gear ring and the small gear ring of the speed gear 22 are respectively meshed with the driving gear 21 and the driven gear 23, and the crank 25 is transmission-connected to the upper center of the driven gear 23.

[0080] In a specific embodiment, the power source includes a motor 4 and a power supply 5. The motor 4 provides power output. The power supply 5 uses a high-power lithium battery, which can provide long-term continuous endurance and support the long-term flight of the aircraft. The power supply 5 is located at the lower part of the fuselage to ensure the stability of the center of gravity of the aircraft. The output shaft of the motor 4 is fixedly connected to the driving gear 21. The two large and small gear rings of the speed change gear 22 are respectively engaged with the driving gear 21 and the driven gear 23. The driven gear 23 is connected to the crank 25. The swing shaft is inserted into the middle slide groove of the slider mechanism 16, and the slider mechanism 16 is driven to reciprocate along the guide rod 15 by swinging.

[0081] In a specific embodiment, the wing 1 is the core component of a flapping-wing aircraft. The four wings 1 are distributed in a rectangular shape, with a flapping amplitude of 120°. The "flapping" mechanism can be used on all four sides to increase lift, thereby achieving efficient aerodynamic performance and flexible movement characteristics.

[0082] In some optional embodiments, the wing 1 is made of a material with good flexibility and high strength, such as carbon fiber, PA12 nylon, polyparaxylene (Parylene-C), polyimide film, flexible polyvinyl chloride, and P31N kite cloth.

[0083] like Figures 14 to 17As shown, the attitude control system of the four-flapping-wing aircraft of the present invention includes a plurality of servos, two rotary servos (servo 1 3 and servo 2 13) and two straight servos (servo 3 8 and servo 4 9), which are respectively used to control the pitch and roll attitude of the aircraft, and realize yaw control by the difference in the lateral and longitudinal flapping planes between the front and rear wings. The two rotary servos are responsible for controlling the roll attitude of the aircraft, and the servos change the angles of the left and right wings 1 by controlling the roll frame 14 to control the roll of the aircraft on the horizontal plane. The two straight servos are used to control the pitch attitude of the aircraft. By adjusting the pitch angles of the front and rear pairs of wings 1, the aircraft can perform up and down pitching movements. At the same time, the coupling effect of the pitch difference between the front and rear wings 1 and the rotation attitude control of the left and right wings 1 enables the aircraft to realize yaw control.

[0084] The embodiment of the present invention further provides a four-flapping-wing aircraft attitude control system, based on the four-flapping-wing aircraft described in the above embodiment, comprising:

[0085] Collect attitude data of the four-wing aircraft in real time, and send control signals to the motor 4 and four steering gears according to the preset attitude expected value;

[0086] The lift is controlled by the rotation speed of the motor 4 to achieve the control of the flight altitude;

[0087] The rudder controls the roll frame 14 to change the angles of the left and right wings 1 to control the roll of the aircraft on the horizontal plane;

[0088] The pitching frame 20 is controlled by the steering gear to change the pitch angles of the front and rear pairs of wings 1, so that the aircraft can perform up and down pitching movements;

[0089] Yaw control is achieved by differences in lateral and longitudinal flapping planes between the four wings 1 .

[0090] The working principle of the four-flapping-wing aircraft and its attitude control system of the present invention is as follows:

[0091] The string drive system uses the pulling characteristics of the string 17 to drive the wing 1 to flap. The motor 4 or motor is used as a driving source to generate a rotational motion, and the rotational motion is converted into the reciprocating motion of the slider mechanism 16 through the crank 25. One end of the string 17 is connected to the slider mechanism 16 and connected to the wing root pulley 11. The string 17 is periodically applied with tension through the guide wheel 10, thereby driving the wing 1 to flap. When the motor 4 drives the crank 25 to rotate, the string 17 is tightened, driving the wing 1 to flap quickly, and realizing the rising or falling action.

[0092] During the flapping of the wing 1, the pulling speed of the general string 17-crank 25 will change dramatically, especially when the wing 1 is rapidly accelerated or decelerated, which will generate large mechanical stress. The strings 17 at both ends of the designed slider mechanism 16 are pulled during the reciprocating motion. Since there are four strings 17 at each end of the slide rod connected to the fixed point of the pulley 11 by cross-linking, this design ensures that when the slider mechanism 16 moves in a certain direction, the string 17 at the other end simultaneously applies corresponding tension through the pulley 11 or the fixed point to maintain the tension balance.

[0093] At each stage of the movement of the slider mechanism 16, the four strings 17 are divided into two groups and are located on both sides of the slider. Through cross connection, it is ensured that the slider is always subjected to symmetrical force distribution during the pulling process. This can effectively prevent the slider from tilting, shaking or uneven force during reciprocating motion. Buffering is provided at key points, thereby extending the life of the system and reducing mechanical losses.

[0094] The pulley 11 or guide wheel 10 optimizes the transmission path. The pulley 11 system helps guide the movement path of the string 17, ensuring a smooth transmission process, reducing the friction and loss of the string 17, and can also adjust the angle of the string 17 transmission to better adapt to the spatial layout. The design of the slider mechanism 16 realizes balanced and symmetrical force on the slider in reciprocating motion by cross-linking four strings 17 at both ends, greatly improving the system's motion stability and transmission efficiency. Through this design, the slider mechanism 16 moves more smoothly, mechanical wear is reduced, and the structure is simple and easy to implement, which is very suitable for use in lightweight flapping-wing aircraft.

[0095] The two straight-line servos (servo three 8 and servo four 9) are responsible for pitch control and adjust the relative flapping plane angles of the front and rear wings 1. When the angle of the front wing 1 increases and the angle of the rear wing 1 decreases, the front of the aircraft will rise and the rear will fall, achieving an upward flight attitude adjustment, such as Fig.15 shown.

[0096] Two rotary servos (servo 1 3 and servo 2 13) are responsible for roll control, and the two servos adjust the flapping angles of the left and right wings 1. By changing the flapping difference of the left and right wings 1 (for example, one side of the wing 1 is slightly raised and the other side is slightly lowered), left and right roll can be achieved, such as Fig.16 shown.

[0097] Yaw is achieved by coupling the pitch difference between the front and rear wings 1 and the roll attitude, such as Fig.17 shown.

[0098] Compared with the prior art, the embodiments of the present invention at least disclose the following beneficial effects:

[0099] The embodiment of the present invention designs a deflectable wing beam structure at the root of the wing 1, so that the flapping-wing aircraft can passively adjust the angle of attack with the airflow during flight, thereby improving the aerodynamic efficiency of the wing 1, increasing the thrust output, and improving the flexibility and adaptability of the aircraft. Therefore, the four-flapping-wing aircraft of the embodiment of the present invention can improve aerodynamic performance.

[0100] The embodiment of the present invention can alleviate the stress concentration problem in the acceleration and deceleration process of the drive system and the wing 1. The crank 25 connecting rod 7 and the chord drive system of the existing flapping-wing aircraft are prone to sudden stress when the drive system and the wing 1 are rapidly accelerated or decelerated, resulting in mechanical fatigue or damage. The present invention alleviates the stress concentration problem, prolongs the service life of the drive system, and improves the overall reliability by introducing a drive system combining the chord 17 with the crank 25 slider mechanism 16.

[0101] The attitude control methods of traditional flapping-wing aircraft (such as center of gravity change control, wing root twist control, etc.) have problems such as slow response speed and low control accuracy, making it difficult to meet the attitude adjustment requirements during fast flight. The present invention adopts a combination of two servos and two linear servos to control the pitch torque by synchronously changing the inclination of the flapping planes of the front and rear wings 1; to control the roll torque by synchronously changing the inclination of the wings 1 on both sides laterally; and to control the yaw torque by asynchronously changing the flapping planes of the front and rear wings 1 and synchronously changing the inclination of the lateral flapping plane. The four-flapping-wing aircraft and its attitude control system of the present invention simplify the structure, reduce control delay, and significantly improve the attitude control efficiency and accuracy of the aircraft.

[0102] In dealing with complex flight environments, existing flapping-wing aircraft have difficulty adjusting their attitude quickly when dealing with complex environments (such as sudden airflow or narrow space), which affects flight stability. The control system of the present invention is based on the change of the flapping plane angle of the wing 1, has a fast response speed, can efficiently deal with complex flight environments, and can quickly adjust the attitude under the influence of sudden airflow, especially to ensure the stability and safety of the aircraft.

[0103] Although the existing multi-axis control system can achieve attitude adjustment, the system structure is complex, which increases the weight of the aircraft and limits the lightweight design and miniaturization of flapping-wing aircraft. The present invention simplifies the control system by combining a servo with a linear servo, reduces the weight of the aircraft while ensuring efficient attitude control, and is suitable for the design requirements of micro-aircraft.

[0104] The present invention realizes the control of yaw attitude through the coupling effect of the differential pitch and roll attitude of the front and rear winglets. This control method not only reduces the complexity of yaw control, but also realizes efficient lateral attitude adjustment through the synergy of roll and pitch, which is suitable for the needs of high maneuverability flight.

[0105] The details not disclosed in the present invention are conventional technical means well known to those skilled in the art.

[0106] In the description of the present invention, it should be understood that the terms "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside" and "outside" etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the present invention.

[0107] The embodiments described above are only descriptions of the preferred modes of the present invention, and are not intended to limit the scope of the present invention. Without departing from the design spirit of the present invention, various modifications and improvements made to the technical solutions of the present invention by ordinary technicians in this field should all fall within the protection scope determined by the claims of the present invention.

Claims

1. A four-flapping wing aircraft, characterized in that: include: A fixed frame (24); a guide rod (15) is provided in the middle of the fixed frame (24) along the front-rear direction, a slider mechanism (16) is reciprocatingly slid on the guide rod (15), and the fixed frame (24) has four cantilevers extending in the front-rear direction; Two pitch frames (20) are symmetrically rotatably connected to the front and rear ends of the fixed frame (24), and the pitching motion of the two pitch frames (20) is controlled by a straight-line steering gear; A rolling frame (14) is connected to the outer side of the pitch frame (20) in a rotational manner in the front-rear direction, and the rolling frame (14) is controlled to rotate by a rotating steering gear, thereby driving the front and / or rear wings (1) to roll left and right; There are four pulleys (11), the four pulleys (11) are horizontally rotatably connected to the left and right ends of the two pitch frames (20), and a wing (1) is hingedly connected to the outer peripheral side of each pulley (11); A string drive system, comprising a guide wheel (10) rotating on the four cantilevers, a string (17) connecting a slider mechanism (16) and a pulley (11), and a crank (25) driving the slider mechanism (16) to reciprocate; one end of the string (17) is fixed to one side of the slider mechanism (16), and the other end is wound around the guide wheel (10) and connected to the pulley (11); two strings (17) are wound around each pulley (11) in the forward and reverse directions, and the two strings (17) are respectively connected to the front side and the rear side of the slider mechanism (16) through different guide wheels (10); the crank (25) is connected to a power source through a transmission system.

2. The four-flapping-wing aircraft according to claim 1, characterized in that: Wing fixing arms (28) are symmetrically extended on the left and right sides of the pitch frame (20), and the wing fixing arms (28) are rotatably connected to the pulley (11).

3. The four-flapping-wing aircraft according to claim 2, characterized in that: A sliding limiter fixing arm (29) is symmetrically extended on the left and right sides of the rolling frame (14), a sliding limiter (12) is fixedly connected to the sliding limiter fixing arm (29), and the sliding limiter (12) is movably connected to the wing (1) to change the flapping plane of the wing (1).

4. The four-flapping-wing aircraft according to claim 3, characterized in that: A wing support beam (18) and a wing root deflection beam (19) are respectively provided on the upper edge and the side edge of the wing (1), and one end of the wing support beam (18) passes through the sliding limiter (12) and is hinged to the peripheral side of the pulley (11).

5. The four-flapping-wing aircraft according to claim 4, characterized in that: The wing root deflection beam (19) is hinged to the end of the wing fixing arm (28) via hinge structure 2.

6. The four-flapping-wing aircraft according to claim 1, characterized in that: The linear steering gear is fixedly connected to the bottom of the fixed frame (24), and has a movable block on one side facing the pitch frame (20) for reciprocating up and down. The movable block is hinged to a connecting rod (7), and one end of the connecting rod (7) away from the movable block is hinged to the pitch frame (20).

7. The four-flapping-wing aircraft according to claim 6, characterized in that: The rotary servo is fixed in the pitch frame (20), and a rudder cap (6) is fixedly connected to the output shaft of the rotary servo. The rudder cap (6) is an equal-arm lever structure, the center of which is fixedly connected to the output shaft of the rotary servo, and the two ends are hinged to the rolling frame (14) through a connecting shaft.

8. The four-flapping-wing aircraft according to claim 1, characterized in that: The wings (1) are distributed in a rectangular shape, and the flapping amplitude is 120°.

9. The four-flapping-wing aircraft according to claim 1, characterized in that: The power source comprises a motor (4) and a power source (5), and the motor (4) and the power source (5) are fixed at the bottom center of the fixed frame (24).

10. A four-flapping-wing aircraft attitude control system, based on the four-flapping-wing aircraft according to any one of claims 1 to 9, characterized in that: include: The attitude data of the four-wing aircraft are collected in real time, and control signals are sent to a motor (4) and four steering gears according to a preset attitude expectation value; lift is controlled by the rotation speed of the motor (4) to achieve control of flight altitude; the angles of the left and right wings (1) are changed by the steering gear to control the roll frame (14) to control the roll of the aircraft on the horizontal plane; the pitch angles of the front and rear pairs of wings (1) are changed by the steering gear to control the pitch frame (20) to make the aircraft perform up and down pitching movements; and yaw control is achieved by the difference in the lateral and longitudinal flapping planes between the four wings (1).

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