JP2018134909A - Unmanned aircraft - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide an unmanned aircraft that can stabilize flight control in such as transition of flight mode.SOLUTION: A multi-copter 101 includes: horizontal rotor units 130A to 130D providing lifting force to an airframe 110 and stationary blades 120A, 120B; and propulsion rotor units 140A, 140B providing thrust force to the airframe. A gyro device 210 as a counter balance element is provided on an opposite side to the propulsion rotor units relative to a gravity center G of the airframe. The gyro device has a principal axis of inertia parallel to a roll axis of the airframe. Thus, violent fluctuation in pitch angle that tends to be generated in transition to horizontal flight from vertical flight can be suppressed.SELECTED DRAWING: Figure 6

Description

  The present invention relates to an unmanned aerial vehicle of a composite rotary wing type that combines a vertical flight capability with a multi-copter rotor and a horizontal flight capability with a fixed wing of an airplane.

  A multicopter, which is a type of unmanned aerial vehicle, includes a plurality of rotor blades that are driven to rotate about a vertical axis. Such an unmanned aerial vehicle is called a VTOL (vertical take-off and landing) aircraft because it can perform vertical flight, horizontal flight, and the like by controlling the rotational speed of each rotor blade. Because the multicopter has hovering flight capability, details of the damage situation by aerial photography in natural disasters such as earthquakes, tsunamis, volcanoes, and human disasters such as chemical factory explosions and marine accidents, etc. The expectation is given to the achievements such as a survey.

  As such a rotary wing type multicopter, for example, in Patent Document 1, a plurality of inner rotor units are arranged on a first circumference separated by a predetermined distance from the center of the fuselage, and more than a predetermined distance from the center of the fuselage. A rotary airfoil machine in which a plurality of outer rotor units are arranged on a second circumference separated by a large predetermined distance, the body is raised by the operation of the inner rotor unit, and the attitude of the body is controlled by the operation of the outer rotor unit. is suggesting.

JP 2014-240242 A

  In the rotary blade type machine described in Patent Document 1 described above, even if the rotational speed of the rotor blades of the outer rotor unit is lowered, a large moment for tilting the fuselage can be generated with a small lift, and the response time for attitude control is shortened. be able to.

  By the way, a general rotary wing type machine has an ascending rotary wing having a vertical axis, and thus has an excellent hovering capability. Further, when flying horizontally, the forward direction is obtained by inclining the direction of travel of the fuselage so as to be lowered forward. That is, in order to obtain a thrust in the horizontal direction, the rotary wing type machine tilts the fuselage by changing the rotational speed of the front and rear rotary blades, and changes the pitch angle of the rotary blades. In this way, by changing the pitch angle of the rotary wing, the rotary wing type aircraft can fly horizontally, but it is not suitable for long-distance flight and high-speed flight compared to fixed wing type aircraft such as radio controlled airplanes, for example. It is.

  In addition, the battery-driven rotary wing type aircraft is limited in flight time and flight distance according to the storage capacity. In particular, if the size of the aircraft is reduced, the capacity of the battery that can be mounted cannot be increased, and the flight time and flight distance are further shortened.

  In order to make up for the deficiencies in such a rotary wing aircraft, for example, it is conceivable to incorporate a flight capability by a fixed wing of an airplane. In this case, fixed wings can be provided on the left and right sides of the airframe, and propulsion rotor blades can be arranged, for example, behind the airframe. For example, when performing horizontal flight from a hovering state, it is considered that high-speed flight can be achieved by controlling so as to gradually switch the thrust from the ascending rotor blade to the thrust force from the propelling rotor blade.

  However, in flight experiments using actual prototypes that the inventors have conducted so far, it is difficult to control the attitude of the aircraft, especially when shifting from the hovering state (vertical flight mode) to the horizontal flight state (horizontal flight mode). In order to make the system always stable, autonomous control is required to have a higher degree of robustness.

  The present invention has been made in view of such a situation, and an object of the present invention is to provide an unmanned aerial vehicle capable of stabilizing flight control at the time of transition between the vertical flight mode and the horizontal flight mode. .

  In order to solve the above-described problems, the present invention provides an unmanned aerial vehicle including a fixed wing, a vertical take-off and landing unit, and a thrust generation unit, and a counterbalance element is provided on the opposite side of the thrust generation unit with respect to the center of gravity of the fuselage. The counter balance element is a gyro device having an inertia main axis parallel to the roll axis of the airframe.

  In the unmanned aerial vehicle, the counterbalance element may be a horizontal rotary wing.

  In the unmanned aerial vehicle, it is preferable that the thrust generating means includes a rotary wing having a rotation axis parallel to the roll axis of the airframe.

According to the unmanned aircraft of the present invention, the center of gravity can be placed near the center of the aircraft by providing the counter balance element on the opposite side of the thrust generating means with respect to the aircraft center of gravity. In addition, the counter balance element is a gyro device having an inertial main axis parallel to the roll axis of the fuselage, so that severe fluctuations in the pitch angle that tend to occur especially when shifting from vertical flight to horizontal flight are suppressed. Can do. The counter balance element may be a horizontal rotary blade.
For these reasons, the attitude control of the fuselage is more robust than the conventional one, and in particular, the flight control at the time of transition to the flight mode can be stabilized.

1 is an external perspective view of a multicopter according to an embodiment of the present invention. It is a top view of the multicopter of FIG. It is a figure for demonstrating the raise control of the multicopter of FIG. It is a figure for demonstrating the roll control of the multicopter of FIG. It is a figure for demonstrating the yaw control of the multicopter of FIG. It is a figure for demonstrating the control apparatus mounted in the multicopter of FIG. It is a lift distribution diagram for demonstrating the autonomous control by the control apparatus of FIG. It is a top view of the multicopter by other embodiments of the present invention. It is a top view of the multicopter by further another embodiment of the present invention.

  Hereinafter, an embodiment in which the present invention is applied to a multicopter including a plurality of rotor blades as an unmanned aerial vehicle according to the present invention will be described with reference to the drawings. First, the configuration of the multicopter will be described with reference to FIGS. 1 and 2. FIG. 1 is an external perspective view of a multicopter 100 according to an embodiment of the present invention, and FIG. 2 is a plan view showing the multicopter 100.

  As shown in these drawings, the multicopter 100 includes fixed wings 120 </ b> A and 120 </ b> B on the left and right of the fuselage 110. In addition, you may equip the edge part of fixed wing | blade 120A, 120B with turbulent wing | blade 121A, 121B for making flight more stable.

  Four circular through-holes 111 are formed in the body 110. Note that the number of through-holes 111 is not limited to four, and may be six or more. These through-holes 111 are formed along a circumferential direction centered on the center of gravity G of the body 110 (more precisely, the center of gravity including the body 110 and the fixed wings 120A and 120B).

  Further, horizontal rotor units 130 </ b> A to 130 </ b> D, which are vertical takeoff and landing means, are disposed in the respective through holes 111. In the present embodiment, four horizontal rotor units 130 </ b> A to 130 </ b> D are arranged at symmetrical positions around the center of gravity G of the airframe 110, but may be six or more.

  Each horizontal rotor unit 130 </ b> A to 130 </ b> D is fixed to a support frame 112 spanned in the through hole 111. Each of the horizontal rotor units 130 </ b> A to 130 </ b> D includes a vertical axis motor 131 and a rotary blade 133 attached to the motor axis 132 of the vertical axis motor 131. As the vertical axis motor 131, for example, a servo motor can be used. Here, the rotor blades 133 and 133 of the horizontal rotor units 130A and 130B adjacent to each other are attached in directions opposite to each other. Similarly, the rotor blades 133 and 133 of the horizontal rotor units 130C and 130D are also attached in opposite directions. For example, the rotor blades 133 and 133 of the horizontal rotor units 130A and 130D that are in a symmetrical positional relationship with respect to the center of gravity G are attached in the same direction.

  Although details will be described later, the rotor blades 133 and 133 of the horizontal rotor units 130A and 130D rotate, for example, in a clockwise direction (CW: clockwise), and the rotor blades 133 and 133 of the horizontal rotor units 130B and 130D For example, it is driven to rotate in a counterclockwise direction (CCW: counterclockwise). In the present specification, rotation in the clockwise direction (CW) is defined as normal rotation, and rotation in the counterclockwise direction (CCW) is defined as reverse rotation.

  Further, by controlling the rotational speed of the rotary blades 133 of the horizontal rotor units 130A to 130D, the postures around the roll axis, the pitch axis, and the yaw axis are corrected as well as rising and lowering.

  Note that the horizontal rotor units 130 </ b> A to 130 </ b> D are not limited to the center in the through-hole 111, but may be suspended below the airframe 110 or may be arranged to protrude above the airframe 110.

  As a thrust generating means of the multicopter 100, two propulsion rotor units 140A and 140B are provided at the rear of the fuselage 110. Each propulsion rotor unit 140 </ b> A, 140 </ b> B includes a horizontal axis motor 141 and a rotary blade 143 attached to the motor shaft 142 of the horizontal axis motor 141. The motor shaft 142 to which the rotary blade 143 is attached is provided in the same direction (parallel) as the roll shaft of the machine body 110.

  For example, a servo motor can be used as the horizontal axis motor 141 of the propulsion rotor units 140A and 140B. In this embodiment, two propulsion rotor units 140A and 140B are arranged. However, the present invention is not limited to this example, and there may be one or three or more. Further, the horizontal axis motor 141 of the propulsion rotor units 140A and 140B and the vertical axis motor 131 of the horizontal rotor units 130A to 130D may have the same specification or different specifications. Good.

  A control device 150 that controls the horizontal rotor units 130 </ b> A to 130 </ b> D and the propulsion rotor units 140 </ b> A and 140 </ b> B is provided below the body 110. Details of the control device 150 will be described later.

  Here, the horizontal rotor units 130 </ b> A to 130 </ b> D and the propulsion rotor units 140 </ b> A and 140 </ b> B are equally arranged on the left and right along the front-rear center line (roll axis) of the body 110. Further, the horizontal rotor units 130A to 130D and the propulsion rotor units 140A and 140B are arranged so that the position of the center of gravity G of the body 110 (precisely, the center of gravity including the body 110 and the fixed wings 120A and 120B) does not shift. 110. Further, the control device 150 is also arranged at the approximate center of the airframe 110 so that the position of the center of gravity G of the airframe 110 does not shift. As described above, the horizontal rotor units 130A to 130D, the propulsion rotor units 140A and 140B, and the control device 150 are arranged at appropriate positions on the fuselage 110 so that the position of the center of gravity G of the fuselage 110 does not shift. 100 horizontal posture control can be stabilized.

  Next, an outline of operation control such as ascending by the horizontal rotor units 130A to 130D will be described with reference to FIGS. 3A, 3B, and 3C. 3A is a diagram for explaining the ascent control, FIG. 3B is a diagram for explaining the roll control around the roll axis, and FIG. 3C explains the yaw control (direction control) around the yaw axis. FIG.

  As described above, the rotor blades 133 of the horizontal rotor unit 130A and the horizontal rotor unit 130D rotate forward (for example, CW), and the rotor blades 133 of the horizontal rotor unit 130B and the horizontal rotor unit 130C rotate in reverse (for example, CCW). Is done. Thereby, the reaction caused by the rotation of the adjacent rotor blades 133 is canceled, and the attitude of the multicopter 100 is stabilized. Further, the rotating blades 133 of the horizontal rotor units 130 </ b> A to 130 </ b> D rotate at the same time, so that the rising posture is stabilized by the gyro effect.

  Here, as shown in FIG. 3A, when the ascent control is performed, the rotational speed of each rotor blade 133 is controlled to be a predetermined value (for example, a command value of a program). At this time, lift by the rotating blade 133 is generated in the fuselage 110 in the direction of the arrow, and when the lift exceeds the weight of the fuselage 110, the multicopter 100 is raised.

  Next, as shown in FIG. 3B, when roll control is performed, the rotational speed of the rotary blades 133 of the horizontal rotor units 130A to 130D is changed. That is, the rotational speed on the horizontal rotor units 130B and 130D side is controlled to be higher than the rotational speed on the horizontal rotor units 130A and 130C side. As a result, the lift by the rotor blades 133 on the horizontal rotor units 130B and 130D side becomes higher than the lift by the rotor blades 133 on the horizontal rotor units 130A and 130C, and the fuselage 110 of the multicopter 100 tilts in the direction of the arrow.

  Next, as shown in FIG. 3C, when performing yaw control, the rotational speed of the rotary blades 133 of the horizontal rotor units 130A to 130D is changed. That is, the rotational speed of the rotary blades 133 of the horizontal rotor units 130A and 130D is controlled to be higher than the rotational speed of the rotary blades 133 of the horizontal rotor units 130B and 130C. At this time, the reaction caused by the rotation of the rotating blade 133 rotating forward in the horizontal rotor units 130A and 130D becomes larger than the reaction caused by the rotation of the rotating blade 133 rotating in the horizontal rotor units 130B and 130C. Thereby, the body 110 of the multicopter 100 changes its direction in the direction of the arrow.

  As is well known, a general multicopter obtains a forward thrust by, for example, tilting the advancing direction side of the aircraft so as to be lowered forward when flying horizontally. However, the multicopter 100 according to the present embodiment is configured such that thrust is obtained by the propulsion rotor units 140A and 140B as described below, and the horizontal posture is maintained even during horizontal flight. As a result, the second lift FL2 generated by the fixed wings 120A and 120B is maximized, and the influence on the control due to a disturbance such as wind is minimized.

  Next, the configuration of the control device 150 will be described with reference to FIG. The control device 150 simultaneously controls the horizontal rotor units 130A to 130D that are vertical takeoff and landing means and the propulsion rotor units 140A and 140B that are thrust generation means.

  More specifically, the control device 150 includes a vertical flight controller (VFC) 163 that controls the number of rotations of each vertical axis motor 131 of the horizontal rotor units 130A to 130D, and each of the propulsion rotor units 140A and 140B. A horizontal flight controller (HFC) 162 that controls the number of rotations of the vertical axis motor 131, and a general control unit 161 that controls the cooperative control of the vertical flight control unit 162 and the horizontal flight control unit 163 are provided. .

  The control device 150 also receives detection signals from various sensors such as a GPS (global positioning system) sensor 155, an airspeed sensor 151, and a gyro sensor 153. Although not shown, the control device 150 may receive a signal from an atmospheric pressure sensor for estimating the altitude of the airframe 110. In addition, information signals from various measuring devices such as a visible camera, an infrared camera, a γ-dose measuring device, and a gas detector attached to the airframe 110 can be input in accordance with the purpose of investigation by the flight of the multicopter 100. Further, a transmitter that wirelessly transmits information measured by these devices to the base device, a receiver that receives a steering signal for remotely controlling the multicopter 100, and the like may be provided.

  The vertical flight control unit 162 uses the horizontal rotor units 130A to 130D, which are vertical takeoff and landing means, in accordance with the turning (yaw) angular velocity command signal and lift command signal output from the overall control unit 161, and ascends, descends, and hovers the airframe 110. Autonomous horizontal posture control including turning. The control amount signal for generating the first lift FL1 output from the vertical flight control unit 162 is amplified by the horizontal rotor drive unit 164, thereby controlling the vertical axis motors 131 of the horizontal rotor units 130A to 130D. It is driven at a different rotational speed.

The vertical flight control unit 162 can autonomously control the multicopter 100 based on the detection signal from the gyro sensor 153. The gyro sensor 153 is, for example, a six-axis gyro sensor. In this case, the angular velocity ω _R around the roll axis and the acceleration α _R around the roll axis, the angular velocity ω _P around the pitch axis, and the acceleration α _P around the pitch axis, Detection signals of the angular velocity ω _Y around the yaw axis and the acceleration α _{Y in} the yaw axis direction are input to the vertical flight control unit 162.

  Next, according to the thrust command signal output from the overall control unit 161, the horizontal flight control unit 163 performs control for causing the airframe 110 to fly forward in the roll axis direction using the propulsion rotor units 140A and 140B as thrust generation means. Do. The control amount signal output by the horizontal flight control unit 163 is amplified by the propulsion rotor drive unit 165, and the horizontal axis motors 141 of the propulsion rotor units 140A and 140B are driven at the controlled rotational speed. For example, the navigator 170 outputs a thrust command signal calculated according to a flight plan registered in advance to the horizontal flight control unit 163, and feeds back a ground speed GS (ground speed) obtained from the GPS sensor 155 to the horizontal flight control unit 163. be able to.

  In the present embodiment, during the flight of the multicopter 100, the first lift LF1 generated by the rotation of the rotary blades 133 of the horizontal rotor units 130A to 130D and the second lift LF2 generated by a difference in atmospheric pressure between the fixed wings 120A and 120B. The rotational speed of the rotary blades 133 of the horizontal rotor units 130A to 130D and the rotational speed of the rotary blades 143 of the propulsion rotor units 140A and 140B are set so that the total of Control at the same time.

Details of this control operation by the overall control unit 161 will be described as follows.
First, the navigator 170 outputs a vertical speed command signal calculated according to a flight plan registered in advance to the altitude command control unit 168. The altitude command control unit 168 calculates a target value of the total lift of the aircraft 100 that should obtain the vertical speed commanded by the navigator 170.
The navigator 170 outputs a turn command signal calculated according to the flight plan to the turn command control unit 167. The turn command control unit 167 outputs a turn (yaw) angular velocity command signal to the vertical flight control unit 162 so that the nose is directed in the direction commanded by the navigator 170.

  The airspeed sensor 151 detects the airspeed AS (airspeed) of the multicopter 100 in the atmosphere based on, for example, the difference between the dynamic pressure applied to the airframe 110 and the static pressure around the airspeed sensor 151. The lift conversion element 169 calculates the second lift LF2 generated by the fixed wings 120A and 120B based on the air speed AS obtained from the air speed sensor 151. In addition, in the lift conversion element 169, in order to calculate the second lift LF2, a nonlinear function described in advance using the airspeed AS, the air density, the total area of the fixed blades 120A and 120B, and the lift coefficient is used.

  The altitude command control unit 168 subtracts the value of the second lift LF2 generated by the fixed wings 120A and 120B from the target value of the total lift that raises or lowers the fuselage 100 at the vertical speed commanded by the navigator 170. Perform feedback control. That is, the altitude command control unit 168 outputs a lift command signal to the vertical flight control unit 162 as a command value for obtaining the first lift LF1 generated by the horizontal rotor units 130A to 130D using the subtracted value. Thereby, the first lift LF1 that complements the second lift LF2 by the fixed wings 120A and 120B is generated by the horizontal rotor units 130A to 130D.

The altitude command control unit 168 may input the acceleration α _Y in the yaw axis direction obtained from the gyro sensor 153 and correct the command value of the lift command signal for obtaining the first lift LF1 as needed.

As described above, in the multicopter 100 according to the present embodiment, the vertical flight control unit 162 and the horizontal flight control unit 163 cooperate with each other to simultaneously control the outputs of the horizontal rotor units 130A to 130D and the propulsion rotor units 140A and 140B. Thus, the horizontal posture of the airframe 110 can be maintained and the flight can be stabilized.
Further, in addition to the first lift by the horizontal rotor units 130A to 130D, the second lift LF2 can be obtained by the fixed wings 120A and 120B, and the power consumed by the horizontal rotor units 130A to 130D can be saved. Further, during horizontal flight, the fixed wings 120A and 120B can glide and fly. From these things, flight time and flight distance can be extended significantly compared with the past.

  Next, an example of autonomous control by the control device 150 will be described with reference to the lift distribution diagram of FIG. In addition, the lift distribution diagram of FIG. 5 has shown the distribution of lift in the ideal case in a windless state. Moreover, the magnitude of the lift force in FIG. 5 is indicated by the size of the area indicated by oblique lines in the flight mode transition section, for example. The same applies to the magnitude of lift in the ascending section, the flying section, and the descending section.

  First, when the multicopter 100 is in the landing state indicated by the arrow a and receives a signal for commanding takeoff, the overall control unit 161 gives an operation command to the vertical flight control unit 162 and the horizontal flight control unit 163. .

  At this time, the vertical flight control unit 162 rotates the rotor blades 133 of the horizontal rotor units 130 </ b> A to 130 </ b> D at the rotation speed according to the command value of the lift command signal. At the same time, the horizontal flight control unit 163 rotates the rotor blades 143 of the propulsion rotor units 140A and 140B at a rotation speed according to the command value of the thrust command signal.

  As a result, the multicopter 100 rises while moving forward to a predetermined altitude between time t0 and time t2, and shifts to horizontal flight. At this time, as can be seen from the lift distribution diagram, as the horizontal flight speed increases, the second lift FL2 by the fixed wings 120A and 120B increases, and accordingly, the first lift FL1 by the horizontal rotor units 130A to 130D can be reduced. Become. This means that the control device 150 sets the horizontal rotor units 130A to 130D and the horizontal rotor units 130A to 130D so that the sum of the first lift FL1 by the horizontal rotor units 130A to 130D and the second lift FL2 by the fixed blades 120A and 120B becomes a target value. This is realized by smoothly controlling the outputs of the propulsion rotor units 140A and 140B.

Further, the multicopter 100 gradually decreases the first lift FL1 by the horizontal rotor units 130A to 130D so that the overall lift becomes the target value as described above when shifting from hovering to horizontal flight indicated by an arrow c. However, the output balance of the horizontal rotor units 130A to 130D and the propulsion rotor units 140A and 140B may be smoothly controlled so that the second lift FL2 by the fixed blades 120A and 120B gradually increases.
In this way, the posture of the fuselage 110 can be stabilized when shifting from the vertical flight mode to the horizontal flight mode.

  In addition, when the multicopter 100 performs horizontal flight indicated by an arrow d between time t2 and time t3, the vertical flight control unit 162 uses the horizontal rotor units 130A to 130D to perform horizontal attitude control of the airframe 110, and yaw. Only control (turning control) can be performed. In that case, the multicopter 100 can glide with the lift of the fixed wings 120A and 120B, can fly at high speed, and can suppress the power consumption of the battery, so that the cruising distance can be increased.

  In addition, when the multicopter 100 is in a horizontal flight state indicated by an arrow d and receives a signal for instructing hovering, the overall control unit 161 gives an operation command to the vertical flight control unit 162 and the horizontal flight control unit 163. .

At this time, the horizontal flight control unit 163 gradually decreases the rotational speed of the rotor blades 143 of the propulsion rotor units 140A and 140B from the horizontal flight state indicated by the arrow e at time t3, thereby reducing the flight speed. The second lift FL2 due to the fixed wings 120A and 120B also decreases due to the decrease in the flight speed, but the vertical flight control unit 162 decreases as the rotational speed of the rotary blades 133 of the horizontal rotor units 130A to 130D gradually increases from time t3. Can be supplemented. As the first lift FL1 by the horizontal rotor units 130A to 130D exceeds the second lift FL2 by the fixed wings 120A and 120B, the multicopter 100 can perform stable hovering as indicated by an arrow f.
In this way, the posture of the fuselage 110 can be stabilized even when shifting from the horizontal flight mode to the vertical flight mode.

And even when the vertical flight control unit 162 lands from the hovering state, the fixed wings 120A and 120B are used while gradually increasing the first lift FL1 by the horizontal rotor units 130A to 130D so that the total lift becomes the target value. The output balance of the horizontal rotor units 130A to 130D and the propulsion rotor units 140A and 140B may be smoothly controlled so that the second lift FL2 gradually decreases.
Thereby, between the time t3 of the hovering state and the time t4, the altitude of the multicopter 100 is gradually lowered and can be safely landed as indicated by an arrow g.

FIG. 6 is a plan view of a multicopter 101 according to another embodiment of the present invention. The multicopter 101 according to this embodiment is provided with fixed wings 120A and 120B on the left and right sides of the fuselage 110, and propulsion rotor units 140A and 140B at the rear of the fuselage 110, and opposite to the propulsion rotor units 140A and 140B with respect to the center of gravity G of the fuselage. That is, at the front of the fuselage 110,
A gyro device 210 is provided. Since the gyro device 210 is a counter balance element of the propulsion rotor units 140 </ b> A and 140 </ b> B, the center of gravity G can be placed near the center of the fuselage 110.

  Further, the gyro device 210 is installed with its inertial main axis parallel to the roll axis of the fuselage 110. As a result, it is possible to suppress a considerable amount of fluctuations in the pitch angle that tend to occur particularly when shifting from the vertical flight mode to the horizontal flight mode.

  Further, as shown in FIG. 7, the counterbalance element of the propulsion rotor units 140 </ b> A and 140 </ b> B may be a horizontal rotary blade 220. That is, the horizontal rotor blade 220 whose rotational speed and direction of rotation are controlled by the control device 150 may be provided on the opposite side of the propulsion rotor units 140A and 140B with respect to the body center of gravity G, that is, on the front portion of the body 110.

  The control device 150 can rotate the horizontal rotary blade 220 so that the wind pressure is generated in a direction that suppresses the variation in the pitch angle of the airframe 110. As a result, it is possible to suppress a considerable amount of fluctuations in the pitch angle, which has been apt to occur when shifting from the vertical flight mode to the horizontal flight mode.

  According to the embodiment of FIGS. 6 and 7, the attitude control of the fuselage 110 is more robust than the conventional one, and in particular, the flight control at the time of transition to the flight mode can be stabilized.

100, 101, 102 Multicopter 110 Airframe 111 Through-hole 120A, 120B Fixed wing 130A-130D Horizontal rotor unit 133 Rotor blade 140A, 140B Propulsion rotor unit 143 Rotor blade 150 Control device 151 Airspeed sensor 153 Gyro sensor 155 GPS sensor 161 General control unit 162 Vertical flight control unit 163 Horizontal flight control unit 164 Horizontal rotor drive unit 165 Propulsion rotor drive unit 167 Turn command control unit 168 Altitude command control unit 169 Lift conversion element 170 Navigator 210 Gyro device 220 Horizontal rotor

Claims (3)

An unmanned aerial vehicle including a fixed wing, a vertical take-off and landing means, and a thrust generating means,
A counter balance element is provided on the opposite side of the thrust generating means with respect to the center of gravity of the aircraft,
An unmanned aerial vehicle characterized in that the counterbalance element is a gyro device having an inertial main axis parallel to a roll axis of a fuselage. An unmanned aerial vehicle including a fixed wing, a vertical take-off and landing means, and a thrust generating means,
A counter balance element is provided on the opposite side of the thrust generating means with respect to the center of gravity of the aircraft,
The unmanned aerial vehicle, wherein the counterbalance element is a horizontal rotary wing. The unmanned aerial vehicle according to claim 1, wherein the thrust generation unit includes a rotor blade having a rotation axis parallel to a roll axis of the fuselage.

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