CN221794920U - Tailstock type unmanned fixed wing VTOL aircraft with built-in rotor wing - Google Patents

Abstract

The utility model discloses a tail-seat unmanned fixed-wing VTOL aircraft with built-in rotors, comprising a fuselage, a built-in rotor fixed wing, a wingtip rotor-wing fused ducted power device, and a vertical tail; the built-in rotor fixed wing is composed of a fixed wing and a built-in rotor; the fixed wing is fixedly connected to the left and right sides of the fuselage; the built-in rotor is embedded in the reserved hole of a flat plate extending backward from the trailing edge of the fixed wing and is evenly arranged by supporting the lower support ribs; the vertical tail is located at the tail of the fuselage, a rudder is installed on the rear side of the vertical tail, and the wingtip rotor-wing fused ducted power device is installed at the wingtip of the fixed wing. The utility model improves the stability of the aircraft and optimizes the tilting transition process by arranging the built-in rotor fixed wing and the wingtip rotor-wing fused ducted power device, and at the same time has the characteristics of vertical take-off and landing and high-speed endurance, and improves the aerodynamic performance of the whole machine and increases the endurance mileage of the whole machine.

Description

A tail-mounted unmanned fixed-wing VTOL aircraft with built-in rotors

Technical Field

The utility model relates to a tail-seat type unmanned fixed-wing VTOL aircraft with built-in rotors in the wings, belonging to the field of overall aircraft design.

Background Art

Traditional fixed-wing aircraft have high flight speed and high cruising efficiency, but they rely on corresponding ground facilities, need runways for takeoff and landing, and cannot hover in the air; traditional rotorcraft can take off and land vertically and can hover in the air, but they do not have the inherent advantages of fixed wings, have slow flight speed, low cruising efficiency, and short range. Therefore, it is necessary to combine rotors with fixed wings to achieve complementary advantages. Most of the existing rotor-fixed wing fusion aircraft have poor cruising performance and insufficient aerodynamic optimization, especially in the transition process from vertical takeoff and landing to level flight. In addition, the stability is poor when encountering crosswinds during vertical takeoff and landing, the ability to resist crosswinds is weak, and the overall design is not novel enough.

Therefore, it is necessary to study an unmanned fixed-wing VTOL aircraft with built-in rotor tail to solve the above problems.

Summary of the invention

The utility model provides a tail-seat unmanned fixed-wing VTOL aircraft with built-in rotors in the wings. By arranging a built-in rotor fixed wing and a wingtip propeller fusion ducted power device, the stability of the aircraft is improved, the tilt transition process is optimized, and the aircraft has the characteristics of vertical take-off and landing and high-speed endurance, and the aerodynamic performance of the whole aircraft is improved, thereby increasing the endurance mileage of the whole aircraft.

The technical solution of the utility model is:

A tail-seat unmanned fixed-wing VTOL aircraft with a built-in rotor in the wing comprises a fuselage, a built-in rotor fixed wing, a wingtip rotor-blade fused ducted power device, and a vertical tail; the built-in rotor fixed wing consists of a fixed wing and a built-in rotor; the fixed wing is fixedly connected to the left and right sides of the fuselage; the built-in rotor is embedded in a reserved hole of a flat plate extending backward from the trailing edge of the fixed wing and is evenly arranged by supporting a supporting rib below; the vertical tail is located at the tail of the fuselage, a rudder is installed on the rear side of the vertical tail, and the wingtip rotor-blade fused ducted power device is installed at the wingtip of the fixed wing.

The wingtip-rotor-blade fused ducted power device includes a propeller, a motor compartment, a winglet, a duct, a support frame, and an aileron; the motor compartment is installed at the wingtip of the fixed wing, the propeller is installed at the rear end of the motor compartment, and the winglet is installed at the outer end of the motor compartment; one end of the winglet is fixedly connected to the motor compartment, and the other end of the winglet is fixedly connected to the duct, and the duct is coaxial with the propeller, and the support frame extends backward along the tail end of the duct casing, and the aileron is installed via the support frame.

The support frame is H-shaped: one end is attached to the duct shell and connected to the fixed wing as a reinforcing rib at the connection, the other end is connected to the duct shell, and the other end is a free end, and the rear end of the cross bar is used to install the aileron.

It also includes the landing gear mounted on the lower part of the fuselage.

The beneficial effects of the utility model are:

1. The center of the wingtip propeller-wing fusion ducted power device of the utility model is fixed on the wingtip of the fixed wing, realizing the fusion of seven parts including the propeller, motor compartment, winglet, reinforcing rib, horizontal wing, duct and aileron; fully considering the aerodynamic optimization of propeller-wing fusion and the reasonable allocation of the positions of various aerodynamic components, it can meet the stability and maneuverability of the aircraft in various flight states and greatly improve the aerodynamic performance of the aircraft; the rotation direction of the propeller is opposite to the rotation direction of the fixed wing wingtip vortex, which can resist the wingtip vortex, and at the same time, the propeller wake is made uniform and vortex-free through the internal propeller-wing fusion characteristics, and the wake interacts with the aileron, so that the aileron can better exert its control characteristics under the action of the ideal wake.

2. When the aircraft of the utility model tilts, the force generated by the built-in rotor will form a nose-down moment around the horizontal axis of the aircraft to tilt the aircraft as a whole. At the same time, the ailerons behind the propeller can also make full use of the propeller wake to assist the tilting, thereby improving the tilting efficiency and controllability of the aircraft and making the tilting process safer.

3. The utility model realizes the pitching motion of the aircraft by controlling the rotation speed of the built-in rotor and the deflection of the aileron control surface. Therefore, there is no elevator on the tail of the aircraft, which simplifies the structure of the aircraft. Most importantly, it ensures that the aircraft has sufficient pitching control performance even in the low-speed mode. In addition, the built-in rotor can also suck the gas in the trailing edge vortex area on the upper surface of the wing, reduce the vortex area, thereby inhibiting flow separation and improving the aerodynamic characteristics of the aircraft wing.

4. The utility model fully utilizes the propeller-wing fusion characteristics through the rear-positioned propeller and the winglet design in the duct, which can effectively improve the pulling force generated by the propeller. The design of the winglet can rectify the airflow inhaled by the propeller, making the propeller wake non-rotational and uniform, which is beneficial to improving the maneuverability of the aileron and improving flight efficiency. The utility model optimizes the aerodynamic characteristics of the coupling between the power propeller and the wing through ingenious design, and realizes the fusion of the wing and the rotor more efficiently.

5. The overall design of the utility model takes into account the aerodynamic optimization of the propeller wing and the reasonable allocation of the positions of various aerodynamic components, optimizes the flight process, and makes the aircraft have good stability and maneuverability.

6. In the vertical mode process of the aircraft of the utility model, the built-in rotor can adjust the force and torque according to the crosswind to which the aircraft is subjected, so that the aircraft can still maintain good stability when being disturbed by crosswind during vertical take-off and landing or hovering.

7. The angle between the fixed wing and the longitudinal axis of the fuselage of the utility model is always a favorable angle of attack for the fixed wing. When the aircraft is in the three modes of flight from vertical to level flight, level flight, and level flight to vertical flight, the longitudinal axis of the aircraft fuselage and the direction of the relative incoming flow are always kept the same, and the fixed wing of the aircraft is always kept at a favorable angle of attack when flying.

In summary, based on the ingenious structural design of the utility model, the aircraft of the utility model can achieve vertical take-off and landing and high-speed cruising at the same time. The overall aerodynamic characteristics of the aircraft are good, the lift-to-drag ratio is high, the flight process is smooth, and the transition process from vertical to level flight is highly stable. Compared with traditional VTOL aircraft, it has better hovering and level flight efficiency, and has better hovering stability against crosswinds. The product can be used in military aviation and general aviation industries, and its uses include logistics and transportation, aerial inspections, agricultural and forestry plant protection, fire rescue, post-disaster rescue, military reconnaissance, border patrols, etc.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG1 is a schematic diagram of the three-dimensional structure of the utility model;

Figure 2 is a partial schematic diagram of the utility model;

Fig. 3 is a front view of the utility model;

Figure 4 is a top view of the utility model;

Figure 5 is a right side view of the utility model;

Figure 6 is a schematic diagram of the vertical take-off and landing stage of the utility model;

FIG7 is a schematic diagram of the utility model in the fuselage tilting stage;

Fig. 8 is a schematic diagram of the utility model in level flight stage;

FIG9 is a schematic diagram of the landing stage of the utility model;

Figure 10 is a detailed schematic diagram of the built-in rotor of the wing of the utility model;

Figure 11 is a detailed schematic diagram of the support frame of the utility model;

FIG12 is a front view of the wingtip propeller fusion ducted power device of the utility model;

FIG13 is a rear view of the wingtip propeller fusion ducted power device of the utility model;

FIG14 is a schematic diagram showing the definition of a coordinate system;

The numbers in the figure are: 1-fuselage, 2-fixed wing, 3-wingtip propeller-wing fused ducted power unit, 4-propeller, 5-motor compartment, 6-wingtip winglet, 7-duct, 8-support frame, 9-aileron, 10-internal rotor, 11-vertical tail, 12-rudder.

DETAILED DESCRIPTION

The utility model is further described below in conjunction with the accompanying drawings and embodiments, but the content of the utility model is not limited to the described scope.

Embodiment 1: As shown in Figure 1-14, a tail-seated unmanned fixed-wing VTOL aircraft with built-in rotors in the wings includes a fuselage 1, a built-in rotor fixed wing, a wingtip propeller fusion ducted power unit 3, and a vertical tail 11; the built-in rotor fixed wing is composed of a fixed wing 2 and a built-in rotor 10; the built-in rotor fixed wing is arranged directly below the center of gravity of the aircraft and is fixedly connected to the left and right sides of the fuselage 1; the built-in rotor 10 is embedded in the reserved hole position of the flat plate extending backward from the trailing edge of the fixed wing 2 and is evenly arranged by supporting the support ribs below, and three built-in rotors 10 are embedded in each of the trailing edge extension plates. The wingtip propeller-wing fusion ducted power device 3 includes a propeller 4, a motor compartment 5, a winglet 6, a duct 7, a support frame 8, and an aileron 9; the motor compartment 5 is installed at the wingtip of the fixed wing 2, and the propeller 4 is installed at the rear end of the motor compartment 5, and the winglet 6 is installed at the outer end of the motor compartment 5; one end of the winglet 6 is fixedly connected to the motor compartment 5, and the other end of the winglet 6 is fixedly connected to the duct 7, and the duct 7 is coaxial with the propeller 4, and a support frame 8 is extended backward along the rear end of the duct shell, and the support frame 8 is H-shaped: one side of the end is attached to the duct shell and connected to the fixed wing 2 as a reinforcing rib at the connection, the other side of one end is connected to the duct shell, and the other end is a free end, and the rear end of the crossbar is used to install the aileron 9 (a servo is installed inside the support frame 8 to control the rotation of the aileron), the aileron 9 is located directly behind the duct, and the width of the aileron 9 is flush with the width of the duct output port. The rear position of the aileron 9 can use the propeller wake to perform rolling control on the aircraft. The vertical tail 11 at the tail of the fuselage is a common tail, and a rudder 12 is provided at the rear side of the vertical tail 11 to assist in controlling the yaw movement of the aircraft, and the tail bracket of the vertical tail 11 and the two support frames 8 form a three-point support for the aircraft. The three-point support structure is simple and can enable the aircraft to land stably and stay on the ground stably. Furthermore, a retractable landing gear is installed on the belly of the fuselage, which can be deployed for emergency landing when the power unit of the aircraft fails and the aircraft relies on gliding for forced landing.

Furthermore, the rotation direction of the built-in rotor is: looking up along the vertical axis from the belly of the fuselage, the rotation direction of the left built-in rotor is counterclockwise, and the rotation direction of the right built-in rotor is clockwise. The left and right built-in rotors are positive and negative propellers to each other, and the lift direction is the same as the lift direction of the fixed wing. The opposite rotation directions can make the rotational torques generated by the built-in rotors cancel each other out without providing additional resistance to propeller torque, and can also suppress the fixed-wing wingtip vortex.

For example, the fuselage is 0.55 meters wide and 2 meters long, with an overall streamlined design that can reduce drag during level flight and increase flight endurance; the duct diameter is 0.5 meters, the total weight is 160 kilograms, the fixed wing chord is 1 meter long, the single fixed wing span is 1.1 meters long, the overall maximum width is 3.5 meters, and the angle between the fixed wing chord and the longitudinal axis of the fuselage is a favorable angle of attack for the fixed wing, which can ensure that the fixed wing always remains at a favorable angle of attack during level flight and transition mode, optimize aerodynamic performance, increase lift and reduce drag. The size design can meet the requirement that the maximum width of the aircraft is less than or equal to the national standard one-way lane, and the small aspect ratio design can minimize the width of the aircraft while ensuring lift. The diameter of the reserved holes is 0.16 meters, the distance between adjacent holes is 0.024 meters, the distance between the fuselage and the nearest adjacent reserved hole is 0.052 meters, and the distance between the duct and the nearest adjacent reserved hole is 0.018 meters; a support rib is provided directly below the center of the reserved hole, and the two sides of the support rib are fixed to the outer wall of the duct and the fuselage respectively. The size design satisfies the placement of three built-in rotors on each side, ensuring the utilization rate of the built-in rotors to the greatest extent.

Furthermore, one end of the winglet 6 is fixedly connected to the outside of the motor compartment 5 of the fixed wing tip, and the other end is fixedly connected to the inner wall of the duct 7. Two winglets are installed on each side, and the two are symmetrical about the horizontal plane. They serve as a fulcrum for supporting the duct shell, and can support the duct shell and stabilize the duct while resisting the wingtip vortex. The plane where the winglet is located is at an angle of 60 degrees to the horizontal plane (as shown in Figure 3). It is installed at the front end of the propeller and can rectify the incoming flow, making the propeller wake uniform, non-rotating and more stable. The H-shaped support frame can be used as a reinforcing rib, and can also facilitate the assembly of the aileron 9, and can further serve as a support point for the aircraft.

Furthermore, the built-in rotor is embedded in the reserved hole of the flat plate extending backward from the trailing edge of the fixed wing. The lift direction of the built-in rotor is the same as that of the fixed wing. A total of 6 built-in rotors are installed on both sides, which can provide lift during level flight, provide the torque required for the pitch behavior of the aircraft, increase the ability to resist crosswinds when hovering in the air, and enhance the maneuverability of converting from vertical mode to level flight mode.

Furthermore, the built-in rotor fixed wing is arranged directly below the center of gravity of the aircraft and is fixedly connected to the left and right sides of the fuselage; the thrust center of the wingtip rotor-wing fused ducted power unit is located below the center of gravity of the aircraft; the lift center of the built-in rotor fixed wing is located behind the center of gravity of the aircraft; when the aircraft is flying, the wingtip rotor-wing fused ducted power unit provides a nose-up moment, and the built-in rotor fixed wing provides a nose-down moment. The balance or imbalance of the two moments corresponds to the aircraft behavior of level flight or attitude change.

As shown in FIG. 1 , the wing-built-in rotor tail-seat unmanned fixed-wing VTOL aircraft provided by the utility model is equipped with 6 built-in rotors. In addition, the rolling motion of the unmanned VTOL aircraft of the utility model is realized by the aileron 9, and the rolling moment is generated by controlling the aileron 9 to realize the rolling motion. The yaw of the unmanned VTOL aircraft of the utility model is realized by the rudder 12, and the deflection angle of the rudder 12 is controlled so that the aircraft generates a lateral force to provide a yaw moment to realize the yaw motion. The built-in rotor 10 of the unmanned VTOL aircraft of the utility model can be used to change the pitch angle of the aircraft. When the aircraft is in a balanced state, the pitch angle can be changed by increasing or decreasing the rotation speed of the built-in rotor 10. When the main power propeller 4 fails, the unmanned VTOL aircraft of the utility model can glide with the help of the lift generated by the fixed wing 2, and the pitch motion of the aircraft is controlled by the overall deflection of the aileron 9, and the rolling motion of the aircraft is controlled by the differential deflection of the aileron 9. The yaw motion of the aircraft is controlled by the rudder 12, thereby ensuring the safety and operability of the aircraft.

Furthermore, the unmanned VTOL aircraft of the utility model is bilaterally symmetrical and is a face-symmetrical aircraft. The plane of symmetry is the plane passing through the longitudinal axis and the vertical axis, and the center of gravity of the aircraft is located in the plane of symmetry. The built-in rotor 10 is located behind the center of gravity and far away from the center of gravity, providing a nose-down moment. By increasing the lever arm, the torque generated by the built-in rotor 10 is large enough to balance the nose-up moment generated by the propeller 4 and ensure maneuverability. The rotation axis of the built-in rotor is parallel to the vertical axis, and the aerodynamic direction is upward along the vertical axis. The aerodynamic center of the propeller 4 is located below the center of gravity of the aircraft, and the aerodynamic direction is forward along the longitudinal axis, providing a nose-up moment for the aircraft; the wingtip-blade fusion ducted power device can improve aerodynamic efficiency due to the fusion design of the propeller, motor compartment, winglet, duct, support frame, and aileron, and can reduce the induced drag caused by the wingtip vortex, reduce the damage to lift caused by the bypass flow, and improve the lift-to-drag ratio, so as to achieve the purpose of increasing lift and reducing drag.

The entire process of fuselage tilting is divided into five processes, and its working principle is as follows:

Process 1: vertical take-off process, as shown in Figure 6, in this process, the aircraft stands vertically on the ground, the two support rods and the vertical tail on the aircraft support the aircraft at three points, and the rotation axis of the propeller 4 is parallel to the longitudinal axis. The forces on the aircraft include: the vertical upward force generated by the propeller 4, the horizontal force generated by the built-in rotor 10, the gravity on the aircraft, and the horizontal force generated by the fixed wing 2. The moments on the aircraft include: the head-up moment generated by the propeller 4, the head-down moment generated by the built-in rotor 10, the yaw moment generated by the rudder 12, the head-down moment generated by the aileron 9 as a whole, and the rolling moment generated by the differential of the ailerons 9 on the left and right sides. When the aircraft takes off, the propeller 4 starts to rotate to provide lift, and the lift is greater than the gravity to make the entire aircraft leave the ground. After the aircraft reaches a certain height, the gravity and lift are balanced and it hovers in the air. Since the aircraft is symmetrical on the left and right, the rolling moment and the yaw moment cancel each other out. During the whole process, it is ensured that the nose-up moment generated by the propeller 4 is equal to the nose-down moment generated by the built-in rotor 10 and the aileron 9, that is, the pitch moment balance of the aircraft is ensured. When the aircraft is hovering and is disturbed by crosswinds in any direction, it can be decomposed into vertical upward, downward, left, right, forward, and backward crosswinds. Therefore, only the upward, downward, left, right, forward, and backward anti-crosswind methods are considered for balancing operations. That is, when the aircraft is affected by a vertical backward crosswind, it will have a tendency to move backward. At this time, the rotation speed of the built-in rotor can be increased to make the aircraft produce a downward attitude change. At this time, the propeller pull generates a forward component to resist the front crosswind force and the horizontal component of the built-in rotor. The propeller pull generates an upward component and the built-in rotor pull generates an upward component to offset gravity, and the aircraft maintains torque balance in this state by adjusting the torque. When affected by a vertical forward crosswind, the aircraft will tend to move forward. At this time, the speed of the built-in rotor can be increased. At this time, the propeller pull and gravity are balanced to keep the longitudinal motion, and the built-in rotor pull and crosswind force are balanced to keep the lateral motion. The aircraft can also maintain torque balance in this state by adjusting the torque. When the aircraft is affected by a vertical left crosswind in the hovering state, the aircraft will tend to move to the left. At this time, the ailerons can be adjusted to make the aircraft roll in the corresponding direction, and the left crosswind will be transformed into a backward crosswind, and the balancing operation can be performed. When the aircraft is affected by a vertical right crosswind in the hovering state, the aircraft will tend to move to the right. At this time, the ailerons can be adjusted to make the aircraft roll in the corresponding direction, and the right crosswind will be transformed into a backward crosswind, and the balancing operation can be performed. When resisting the crosswinds on the left and right sides, it is necessary to keep the lift and gravity of the aircraft balanced, and the torque of the aircraft can be balanced by adjusting the torque. When the aircraft is affected by vertically upward and vertically downward crosswinds, gravity and lift are unbalanced, and the hovering state of the aircraft will change. The pulling force can be changed by changing the speed of propeller 4 to achieve balance of the aircraft, and the torque of the aircraft can be balanced by adjusting the torque.

Process 2: Vertical takeoff-level flight transition process, as shown in Figure 7-8, the aircraft starts to enter the tilt transition process from the hovering state. At this time, the forces on the aircraft include: the force parallel to the longitudinal axis generated by the propeller 4, the force parallel to the vertical axis generated by the built-in rotor 10, the gravity of the aircraft, the lift generated by the wings on the left and right sides of the fuselage, and the resistance generated by the fuselage and wings. The moments on the aircraft include: the nose-up moment generated by the propeller 4, the nose-down moment generated by the built-in rotor 10, the adjustable yaw moment generated by the rudder 12, the rolling moment and the adjustable pitch moment generated by the aileron 9, which are convenient for controlling the overall pitching movement of the aircraft. During the tilting process, the built-in rotor 10 increases the rotation speed, generates a nose-down moment, and causes the aircraft to change its posture in a downward pitch. After the tilting begins, the propeller thrust generates a positive component along the longitudinal axis of the ground system, allowing the aircraft to have a forward flying speed along the longitudinal axis. The propeller thrust generates a positive component along the vertical axis of the ground system, allowing the aircraft to have an ascending speed along the vertical axis of the ground system. The sum of the forward flying speed and the ascending speed is the incoming flow speed, and the incoming flow speed is parallel to the longitudinal axis of the fuselage, ensuring that the fixed wing of the aircraft always flies at a favorable angle of attack. During the tilting process, the built-in rotor 10 and the aileron 9 as a whole generate a pitching moment that causes the aircraft to lower its head, causing the aircraft to change its posture in a downward pitch. During the tilting process of the aircraft, since the aircraft is symmetrical on both sides, the rolling moment and the yaw moment on the left and right sides of the aircraft cancel each other out. As the aircraft gradually tilts to a horizontal position as a whole, as shown in Figure 5. The aircraft continuously accelerates during the tilting process to enter the fixed-wing flight mode. During the tilting process of the aircraft, first ensure that the aircraft always maintains a favorable angle of attack of the fixed wing.

Process three: level flight process, as shown in Figure 5, at this time the aircraft is flying forward. During this process, the forces acting on the aircraft include: the forward pulling force generated by the propellers 4 on the left and right sides of the fuselage, the upward lift generated by the wings 2 on the left and right sides of the fuselage, the upward lift generated by the built-in rotor 10, the resistance generated by the aircraft and the gravity acting on the aircraft. The moments acting on the aircraft include: the nose-up moment generated by the propellers 4 on the left and right sides of the fuselage, and the nose-down moment generated by the built-in rotor 10. The yaw moment and roll moment generated by the differential of the propellers on the left and right sides. The aileron 9 can provide a roll moment, and the rudder 12 of the vertical tail 11 can provide a yaw moment.

In the level flight state, the lift of the aircraft is balanced with the gravity, and the thrust is balanced with the resistance. Since the aircraft is symmetrical, the yaw and roll moments on the left and right sides of the aircraft cancel each other out. When the aircraft's nose-up moment is equal to the nose-down moment, the aircraft's pitch moment is balanced. When the aircraft is subject to pitch disturbance, the aircraft's angle of attack changes, and the lift of the fixed wing and the fuselage changes, that is, additional lift is generated. The additional lift will generate a moment in the opposite direction of the disturbance to ensure pitch stability. When the aircraft cannot return to a balanced state by itself, the speed of the propeller 4 and the built-in rotor 10 can also be controlled to adjust the aircraft's pitch moment to ensure pitch stability. Here, the pulling force of the propeller 4 in the longitudinal axis direction can be changed by adjusting the speed change of the propeller 4. At this time, the nose-up and nose-down moments of the front and rear propellers 4 are unbalanced, which will generate a pitch moment that drives the aircraft to pitch, causing the aircraft to pitch. When adjusting the pitch attitude, fine-tuning can be performed through the built-in rotor 10. When the aircraft is subject to rolling disturbance, the lift generated by the wings on the left and right sides of the fuselage is unbalanced, which can generate a stabilizing moment in the opposite direction of the rolling direction. Since the vertical tail 11 is located above and behind the center of gravity, the side force generated will form a stabilizing moment in the opposite direction of the rolling direction on the center of gravity, so that the aircraft maintains rolling stability. When the aircraft cannot return to a balanced state by itself, the aileron 9 can also be controlled to deflect to achieve rolling balance. In addition, by manipulating the deflection of the left and right ailerons, the different lift differences on the two ailerons will generate a moment that drives the aircraft to roll, so that the aircraft performs a rolling motion. The rotation speed difference of the built-in rotors 10 on both sides can also be adjusted to fine-tune the aircraft roll. When the aircraft is subject to heading disturbance, the vertical tail 11 will generate a lateral force in the same direction as the yaw direction under the action of the airflow, so that the aircraft generates a stabilizing moment in the opposite direction of the yaw direction. In addition, when the aircraft is subject to heading disturbance and sideslips, since the aircraft wings can use a sweep angle, the existence of the sweep angle will increase the resistance of the front wing on one side of the sideslip and reduce the resistance on the other side, thereby generating a yaw stabilizing moment. When the aircraft cannot return to a balanced state by itself, the rudder 12 can also be adjusted to ensure yaw stability. In addition, by manipulating the deflection of the rudder 12, a yaw moment that drives the aircraft to yaw can be generated, causing the aircraft to yaw. When the aircraft is affected by wind during level flight, the wind speed direction can be decomposed into the left and right side crosswinds and the wind in the symmetry plane for analysis. As for the effect of the wind in the symmetry plane, due to Therefore, when the aircraft is affected by the wind in the plane of symmetry, the aircraft needs to adjust the angle of attack of the aircraft according to the size and direction of the relative airspeed, so that the aircraft changes its pitch angle, prevents the aircraft from insufficient lift due to a small angle of attack or stalling due to a large angle of attack, and enables the aircraft to maintain a favorable angle of attack. When affected by the wind in the plane of symmetry, since changing the angle of attack of the aircraft requires changing the rotation speed of the built-in rotor 10 of the aircraft, it is necessary to re-ensure the force balance of the aircraft after adjusting the angle of attack when affected by the wind, and achieve torque balance by adjusting the torque. When affected by a vertical left crosswind, the aircraft will tend to move to the left. At this time, the left aileron of the aircraft can be controlled downward and the right aileron can be controlled upward to make the aircraft produce a roll angle to the right. At this time, the built-in rotors on both sides of the aircraft will produce a rightward component force to resist the crosswind. Similarly, when the aircraft is affected by a vertical right crosswind, the aircraft will tend to move to the right. At this time, the left aileron of the aircraft can be controlled to move upward and the right aileron can be controlled to move downward, so that the aircraft will roll to the left. The propeller groups on both sides of the aircraft will generate a left component force to resist the crosswind. When resisting the crosswind on both sides, it is necessary to maintain the balance between the lift and gravity and the thrust and drag of the aircraft, and balance the aircraft torque by adjusting the torque.

Process 4: Level flight-vertical landing transition process, as shown in Figures 7-8, the forces on the aircraft at this time are: the force along the longitudinal axis generated by the propeller 4, the force along the vertical axis generated by the built-in rotor 10, the lift generated by the wing 2 and the gravity of the aircraft. The moments on the aircraft are: the nose-up moment generated by the propeller 4, the nose-down moment generated by the built-in rotor 10, the yaw moment generated by the rudder 12 and the rolling moment generated by the aileron differential. During the tilting process, the speed of the built-in rotor is first reduced to reduce its nose-down moment. The overall nose-down moment is less than the nose-up moment generated by the propeller, and the aircraft raises its head. At this time, the speed of the propeller 4 is continuously reduced to reduce its speed along the longitudinal axis of the fuselage. At the same time, due to the nose-up moment, the longitudinal axis of the aircraft fuselage is facing upward. At this time, the pulling force generated by the propeller 4 is the same as the gravity of the aircraft, achieving the purpose of hovering. During the tilting process, since the aircraft is symmetrical on the left and right, the rolling moment and yaw moment on the left and right sides are equal to achieve rolling and yaw balance.

Process five: Landing process, as shown in Figure 9, after the aircraft tilts to a vertical state as a whole, the propeller 4 still keeps rotating, and the aircraft stabilizes in a hovering state first, and then begins to land slowly. During the landing process, the aircraft is subjected to the following forces: the pulling force along the longitudinal axis generated by the propeller 4, the force along the vertical axis generated by the built-in rotor 10, and the gravity of the aircraft. The propeller speed can be controlled so that the lift of the aircraft is less than the gravity, so that the entire aircraft lands slowly; the moments subjected to the aircraft are: the nose-up moment generated by the propeller 4, the nose-down moment generated by the built-in rotor 10, the yaw moment provided by the anti-torque of the propellers on the left and right sides, and the rolling moment generated by the differential of the left and right ailerons. Since the left and right sides of the aircraft are symmetrical, the rolling moment and yaw moment on the left and right sides of the aircraft cancel each other out. In the whole process, the nose-up moment and the nose-down moment are equal, which can ensure the balance of the pitch moment of the aircraft. The solutions for maintaining the balance of the aircraft after being disturbed by wind in this process are the same as those in process one, and will not be repeated here.

It should be noted that the horizontal axis, the vertical axis, and the vertical axis are specifically defined in the attached figure 14. With the center of gravity of the aircraft as the origin, the X-axis is the positive direction along the longitudinal symmetry plane of the aircraft, which is called the longitudinal axis; the Y-axis is perpendicular to the longitudinal axis in the longitudinal symmetry plane of the aircraft, with the upward direction as the positive direction, which is called the vertical axis; the Z-axis is perpendicular to the longitudinal symmetry plane, with the right direction as the positive direction, which is called the horizontal axis. The positive and negative directions of the rotation of the aircraft around each axis are determined in accordance with the right-hand rule. It should be noted that the words "front", "rear", "left", "right", "up" and "down" used in the following description refer to the directions in the attached figure, where "front" represents the positive direction of the longitudinal axis of the aircraft, "rear" represents the negative direction of the longitudinal axis of the aircraft, "right" represents the positive direction of the horizontal axis of the aircraft, "left" represents the negative direction of the horizontal axis of the aircraft, "up" represents the positive direction of the vertical axis of the aircraft, and "down" represents the negative direction of the vertical axis of the aircraft. The clockwise and counterclockwise directions around each axis are defined by the right-hand rule. The direction that complies with the right-hand rule is "clockwise", and the opposite is "counterclockwise".

The specific implementation modes of the present invention have been described in detail above in conjunction with the accompanying drawings, but the present invention is not limited to the above implementation modes, and various changes can be made within the knowledge scope of ordinary technicians in the field without departing from the purpose of the present invention.

Claims (4)

1. A tailstock type unmanned fixed wing VTOL aircraft with a built-in rotor wing is characterized by comprising a fuselage, a built-in rotor wing fixed wing, a wing tip and oar wing fusion duct power device and a vertical tail wing; the built-in rotor wing fixed wing consists of a fixed wing and a built-in rotor wing; the fixed wings are fixedly connected to the left side and the right side of the machine body; the built-in rotor wing is embedded in reserved hole sites of a flat plate with the rear edge of the fixed wing extending backwards and is uniformly arranged by means of supporting ribs at the lower part, the vertical tail wing is positioned at the tail part of the fuselage, a rudder is arranged at the rear side of the vertical tail wing, and the tip of the fixed wing is provided with a wing tip paddle wing fusion culvert power device.

2. The unmanned fixed wing of tailstock VTOL aircraft with built-in rotor wings according to claim 1, wherein the wing tip-blade fusion ducted power device comprises a propeller, a motor bin, a wing tip winglet, a duct, a support frame, an aileron; the motor bin is arranged at the fixed wing tip, the propeller is arranged at the rear end of the motor bin, and the wing tip winglet is arranged at the outer end of the motor bin; one end of the wing tip winglet is fixedly connected to the motor bin, the other end of the wing tip winglet is fixedly connected to the duct, the duct is coaxial with the propeller, the support frame is extended backwards along the tail end of the duct shell, and the aileron is installed through the support frame.

3. The unmanned fixed wing tailstock VTOL aerial vehicle with built-in rotor of claim 2, wherein the support frame is of the H-type: one end of the cross rod is attached to the duct shell, the cross rod is connected with the fixed wing to serve as a reinforcing rib at the joint, the other end of the cross rod is connected with the duct shell, the other end of the cross rod is a free end, and the rear end of the cross rod is used for installing the aileron.

4. The unmanned fixed wing tail stock VTOL aerial vehicle of claim 1, further comprising landing gear mounted in a lower portion of the fuselage.