Unmanned aerial vehicle

Abstract

This unmanned aerial vehicle is provided with a plurality of rotors, and comprises: an internal combustion engine; a power generation device that is driven by the internal combustion engine to generate power; a battery that stores the power; and a plurality of electrical components to which the power is supplied from the battery. The plurality of electrical components include: a first electrical component group that consumes less power than the threshold power at the time of startup; and a second electrical component group that consumes more power than the threshold power at the time of startup. The second electrical component group includes: a first electrical component that consumes first power at the time of startup; and a second electrical component that consumes second power at the time of startup. After the internal combustion engine is started, the timing at which the first electrical component starts is different from the timing at which the second electrical component starts.

Description

unmanned aerial vehicle

[0001] The present disclosure relates to unmanned aerial vehicles.

[0002] An unmanned aerial vehicle (UAV) is an aircraft that cannot carry a person due to its structure and can fly by remote control or automatic pilot. Rotary-wing unmanned aerial vehicles are unmanned aerial vehicles that obtain lift using propellers that rotate around an axis, i.e., rotors. Small unmanned aerial vehicles equipped with multiple rotors (multi-rotor UAVs) are also called "drones," "multirotors," or "multicopters," and are widely used for applications such as aerial photography, surveying, logistics, and pesticide spraying.

[0003] Patent Document 1 describes an unmanned aerial vehicle (unmanned flying object) that changes its flight position in conjunction with the operation of agricultural machinery.

[0004] Japanese Patent Application Laid-Open No. 2022-104737

[0005] The present disclosure provides a startup sequence for electrical equipment installed on an unmanned aerial vehicle.

[0006] In an exemplary and non-limiting embodiment, the unmanned aerial vehicle of the present disclosure is an unmanned aerial vehicle having multiple rotors, and includes an internal combustion engine, a power generation device driven by the internal combustion engine to generate electricity, a battery that stores the electricity, and multiple electrical components to which power is supplied from the battery, wherein the multiple electrical components include a first group of electrical components that consume less power than a threshold power when started up, and a second group of electrical components that consume more power than the threshold power when started up, and the second group of electrical components includes a first electrical component that consumes a first power when started up, and a second electrical component that consumes a second power when started up, and the timing at which the first electrical components start up after the internal combustion engine starts is different from the timing at which the second electrical components start up.

[0007] According to an embodiment of the present disclosure, a novel startup sequence for electrical components onboard an unmanned aerial vehicle is provided.

[0008] 1 is a block diagram schematically showing several examples of a rotary drive unit that rotates rotors in an unmanned aerial vehicle having multiple rotors. FIG. 2 is a plan view schematically showing one basic configuration example of an unmanned aerial vehicle having multiple rotors. FIG. 3 is a side view schematically showing one basic configuration example of an unmanned aerial vehicle having multiple rotors. FIG. 4 is a plan view schematically showing another basic configuration example of an unmanned aerial vehicle having multiple rotors. FIG. 5 is a block diagram showing an example basic configuration of a battery-powered multicopter. FIG. 6 is a block diagram showing an example basic configuration of a series hybrid multicopter. FIG. 7 is a block diagram showing an example basic configuration of a parallel hybrid multicopter. FIG. 8 is a top view schematically showing a multicopter according to an embodiment of the present disclosure. FIG. 9 is a side view schematically showing a multicopter according to this embodiment. FIG. 10 is a block diagram showing an example system configuration of a multicopter according to this embodiment. FIG. 11 is a block diagram showing an example of an electric power system that supplies electric power from a battery 52 to each electrical component. FIG. 12 is a flowchart illustrating an example of a startup sequence of a first and second electrical component group. FIG. 13 is a block diagram showing an example hardware configuration of a control device according to this embodiment. FIG. 14 is a graph showing the change over time in power consumption of a DC-DC converter, a first electrical component, and a second electrical component.

[0009] An unmanned aerial vehicle with multiple rotors includes a rotary drive unit that rotates the rotors (hereinafter sometimes referred to as "propellers"). Hereinafter, such an unmanned aerial vehicle will be referred to as a "multicopters."

[0010] There are various configurations of the rotary drive device provided in a multicopter. Fig. 1A is a block diagram schematically illustrating four examples of the rotary drive device 3 in the present disclosure.

[0011] The first rotation drive device 3A shown in FIG. 1A has a plurality of electric motors (hereinafter referred to as "motors") 14 that rotate a plurality of rotors 2, and a battery 52 that stores power to be supplied to each motor 14. The battery 52 is, for example, a secondary battery such as a polymer lithium-ion battery. Each rotor 2 is connected to the output shaft of the corresponding motor 14 and rotated by the motor 14. In order to increase the payload and / or flight time, it is necessary to increase the power storage capacity of the battery 52. ​​The power storage capacity of the battery 52 can be increased by increasing the size of the battery 52, but increasing the size of the battery 52 results in an increase in weight.

[0012] The second rotation drive device 3B shown in FIG. 1A includes a power transmission system 23 mechanically connected to the rotor 2 and an internal combustion engine 7a that provides driving force (torque) to the power transmission system 23. The power transmission system 23 includes mechanical components such as gears or belts, and transmits torque from the output shaft of the internal combustion engine 7a to the rotor 2. The internal combustion engine 7a can efficiently generate mechanical energy by burning fuel. Examples of the internal combustion engine 7a include a gasoline engine, a diesel engine, and a hydrogen engine. The number of internal combustion engines 7a included in the rotation drive device 3B is not limited to one.

[0013] The third rotary drive device 3C shown in FIG. 1A includes multiple motors 14, a power buffer 9 that stores power to be supplied to each motor 14, a power generator 8 such as an alternator that generates power, and an internal combustion engine 7a that provides mechanical energy for the power generator 8 to generate electricity. A typical example of the power buffer 9 is a battery such as a secondary battery, but it may also be a capacitor. In the third rotary drive device 3C, even if the power buffer 9 does not have a large storage capacity, the power generator 8 generates power using the driving force (mechanical energy) of the internal combustion engine 7a, thereby enabling an increase in payload and / or flight time. This type of drive is called a "series hybrid drive." The power generator 8 and internal combustion engine 7a in the series hybrid drive are called a "range extender" because they extend the flight distance of the multicopter.

[0014] 1A includes a plurality of motors 14, a power buffer 9 that stores power to be supplied to each motor 14, a power generator 8 such as an alternator that generates power, an internal combustion engine 7a that provides driving force for generating power to the power generator 8, and a power transmission system 23 that transmits the driving force generated by the internal combustion engine 7a to a rotor 2 to rotate the rotor 2. At least one rotor 2 of the plurality of rotors 2 is rotated by the internal combustion engine 7a, and the other rotors 2 are rotated by the motor 14. In the fourth rotary drive device 3D, the mechanical energy generated by the internal combustion engine 7a can also be used to rotate the rotor 2 without being converted into electric power, thereby improving energy utilization efficiency. This type of drive is called a "parallel hybrid drive."

[0015] Fig. 1B is a plan view schematically illustrating one basic configuration example of multicopter 10. The configuration example of Fig. 1B includes the first rotational drive device 3A shown in Fig. 1A as the rotational drive device 3. That is, the rotational drive device 3 (3A) in this example includes a motor 14 and a battery 52. ​​Fig. 1C is a side view schematically illustrating multicopter 10.

[0016] 1B and 1C includes a plurality of rotors 2, an airframe 4, and an airframe frame 5 that supports the rotors 2 and the airframe 4. The airframe frame 5 supports the airframe 4 at its center and rotatably supports the plurality of rotors 2 with a plurality of arms 5A extending outward from the center. A motor 14 that rotates the rotors 2 is provided near the tip of each arm 5A. The airframe 4 and the airframe frame 5 are sometimes collectively referred to as the "airframe 11."

[0017] 1B, the multicopter 10 is a quad-type multicopter (quadcopter) having four rotors 2. The rotors 2 located on one diagonal line rotate in the same direction (clockwise or counterclockwise), while the rotors 2 located on different diagonal lines rotate in opposite directions.

[0018] The main body 4 includes a control device 4a that controls the operation of devices and components mounted on the multicopter 10, a group of sensors 4b connected to the control device 4a, a communication device 4c connected to the control device 4a, and a battery 52.

[0019] The control device 4 a may include, for example, a flight control device such as a flight controller and a host computer (companion computer). The companion computer can perform advanced arithmetic processing such as image processing, obstacle detection, and obstacle avoidance based on the sensor data acquired by the sensor group 4 b.

[0020] The sensor group 4b may include an acceleration sensor, an angular velocity sensor, a geomagnetic sensor, a barometric pressure sensor, an altitude sensor, a temperature sensor, a flow rate sensor, an imaging device, a laser sensor, an ultrasonic sensor, an obstacle contact sensor, and a Global Navigation Satellite System (GNSS) receiver. The acceleration sensor and the angular velocity sensor may be mounted on the airframe main body 4 as components of an IMU (Inertial Measurement Unit). Examples of the laser sensor may include, for example, a laser range finder used to measure the distance to the ground, and a two-dimensional or three-dimensional light detection and ranging (LiDAR).

[0021] The communication device 4c may include a wireless communication module for transmitting and receiving signals via an antenna to a transmitter or ground station (Ground Control Station (GCS)) on the ground, a mobile communication module using a cellular communication network, etc. The communication device 4c may receive signals such as control commands transmitted from the ground and transmit sensor data such as image data acquired by the sensor group 4b as telemetry information. The communication device 4c may have a function for communicating between multicopters and a satellite communication function. The control device 4a can be connected to a computer on the cloud via the communication device 4c. Some or all of the functions of the companion computer may be performed by the computer on the cloud.

[0022] The battery 52 is a secondary battery that stores power by charging and supplies power to the motors 14 by discharging. The battery 52 and the multiple motors 14 operate to rotate the multiple rotors 2, making it possible to generate a desired thrust.

[0023] Each of the multiple rotors 2 generally has multiple blades with a fixed pitch angle and generates thrust by rotation. The pitch angle may be variable. The multiple rotors 2 do not all need to have the same diameter (propeller diameter); one or more rotors 2 may have a larger diameter than the other rotors 2. The thrust (static thrust) generated by a rotating rotor 2 is generally proportional to the cube of the rotor 2 diameter. Therefore, when rotors 2 with different diameters are included, the rotor 2 with a relatively larger diameter may be referred to as the "main rotor," and the rotor 2 with a relatively smaller diameter may be referred to as the "sub-rotor." Note that, regardless of the diameter, the configuration of the rotary drive device 3 may include a rotor 2 capable of generating a relatively larger thrust and a rotor 2 with a relatively smaller thrust. In this case, the rotor 2 capable of generating a relatively larger thrust may be referred to as the "main rotor," and the rotor 2 with a relatively smaller thrust may be referred to as the "sub-rotor." For example, the rotor 2 that generates a relatively large thrust per rotation may be referred to as the "main rotor," and the rotor 2 that generates a relatively small thrust per rotation may be referred to as the "sub-rotor." In one example, the main rotor may be positioned more inward than the sub-rotors. In other words, each rotor 2 may be positioned so that the distance from the center of the airframe to the rotation axis of each main rotor is shorter than the distance from the center of the airframe to the rotation axis of each sub-rotor.

[0024] In this example, the rotary drive device 3 includes a plurality of motors 14. As mentioned above, the rotary drive device 3 may include an internal combustion engine 7a.

[0025] FIG. 1D is a plan view schematically illustrating an example of the basic configuration of a multicopter 10 including a second rotational drive device 3B as the rotational drive device 3. In the example shown in FIG. 1D , an internal combustion engine 7a is supported by the airframe main body 4. In this example, the driving force generated by the internal combustion engine 7a is transmitted to multiple rotors 2 via multiple power transmission systems 23, causing each rotor 2 to rotate. The control device 4a can change the rotational speed of each rotor 2 by controlling each power transmission system 23. The rotational drive device 3B may include a mechanism for changing the pitch angle of each blade of the multiple rotors 2. In this case, the control device 4a may adjust the lift generated by each rotor 2 by controlling the mechanism to change the pitch angle of the blades.

[0026] In a "parallel hybrid drive" in which some of the multiple rotors 2 are rotated by the internal combustion engine 7a and the other rotors 2 are rotated by the motor 14, the internal combustion engine 7a and the battery 52 are supported on the aircraft body 4. At least one rotor 2 of the multiple rotors 2 is connected to the internal combustion engine 7a via the power transmission system 23, and the other rotors 2 are connected to the motor 14.

[0027] In such a parallel hybrid drive, the diameter of one or more rotors 2 rotated by the internal combustion engine 7a may be larger than the diameter of the other rotors 2 rotated by the motor 14. In other words, the internal combustion engine 7a may be used to rotate the main rotor, and the motor 14 may be used to rotate the sub-rotor. In such a case, the main rotor is primarily used to generate thrust, and the sub-rotor is used to generate thrust and for attitude control. The main rotor may also be called a "booster rotor," and the sub-rotor may also be called an "attitude control rotor."

[0028] In the case of a parallel hybrid drive, the internal combustion engine is used for both thrust generation and power generation. By selectively transmitting the driving force (torque) generated by the internal combustion engine to one or both of the rotor and the power generator, it is possible to achieve a good balance between thrust generation and power generation.

[0029] Equipping a multicopter with an internal combustion engine and using it to generate thrust and / or electricity contributes to increased payload and flight time. It is desirable to control the attitude of a multicopter by rotating the propellers with a motor, which has better response characteristics than an internal combustion engine. Therefore, in applications requiring precise control of the multicopter's attitude, it is desirable to employ a parallel hybrid drive or series hybrid drive to increase the payload and flight time. If the rotary drive device 3 is equipped with a mechanism for changing the pitch angle of each blade of the multiple rotors 2, the attitude can also be adjusted by changing the pitch angle of each blade.

[0030] Increased payload and flight time may further expand the applications of multicopters. For example, in the agricultural field, multicopters are currently being used for spraying pesticides or monitoring crop growth conditions. However, by connecting various ground implements (hereinafter, sometimes simply referred to as "implements") to a multicopter, various agricultural tasks can be performed from the air. Agricultural implements are sometimes called "implements." Examples of implements include sprayers that spray pesticides on crops, mowers, seeders, spreaders, rakes, balers, harvesters, plows, harrows, or rotary tillers. Work vehicles such as tractors are not included in the "implements" of this disclosure.

[0031] In the example shown in FIG. 1C , a work implement 200 is coupled to the multicopter 10. The work implement 200 can spray, for example, pesticides or fertilizers on a field or crops within the field. Increasing the payload and flight time allows for a larger and / or more versatile work implement 200. For example, by changing the work implement 200 coupled to the multicopter 10, a variety of ground tasks (agricultural operations) can be performed, including liquid and granular application of pesticides, fertilization, thinning, weeding, transplanting, direct seeding, and harvesting. The work implement 200 may be equipped with a mechanism such as a robotic hand. In this case, a single work implement 200 can perform a variety of ground tasks. If the work implement 200 has a sufficient space to accommodate materials, the work implement 200 can also be used to transport agricultural materials or harvested products over a wide area. The work implement 200 can be coupled to the multicopter 10 in a variety of ways. The multicopter 10 may suspend and tow the work machine 200 by a cable. The work machine 200 towed by the multicopter 10 can also perform ground work while being towed while the multicopter 10 is flying or hovering. The work machine 200 during work may be in the air or on the ground.

[0032] 1C , the multicopter 10 includes a power supply device 76. The power supply device 76 is a device that supplies power to the work machine 200 from a drive energy source, such as the battery 52 or the power generation device 8, included in the multicopter 10. Various functions of the work machine 200 can be performed using this power. The work machine 200 includes actuators such as motors that operate using power obtained from the power supply device 76 of the multicopter 10. The work machine 200 preferably includes a battery that stores power.

[0033] FIG. 2A is a block diagram illustrating an example of the basic configuration of a battery-powered multicopter 10. The battery-powered multicopter 10 includes multiple rotors 12, multiple motors 14 for rotating the rotors 12, multiple ESCs (electric speed controllers) 16 each having a motor drive circuit for driving the motors 14, a battery 52 for supplying power to the corresponding motors 14 via each ESC 16, a control device 4a for controlling the multiple ESCs 16 to control attitude and perform flight, a sensor group 4b, a communication device 4c, and a power supply device 76 electrically connected to the battery 52. ​​For simplicity, FIG. 2A illustrates the rotors 12, motors 14, and ESCs 16 as a single block, but the number of rotors 12, motors 14, and ESCs 16 is actually multiple. This also applies to FIGS. 2B and 2C. The ESC 16 may be included in the control device 4a.

[0034] The control device 4a can receive control commands wirelessly, for example, from a ground station 6 located on the ground via the communication device 4c. The number of ground stations 6 is not limited to one and may be distributed across multiple locations. The communication device 4c can also receive control commands wirelessly from a control device operated by a pilot on the ground. The control device 4a may have the function of automatically or autonomously performing takeoff, flight, obstacle avoidance, and landing operations based on sensor data obtained from the sensor group 4b. The control device 4a may be configured to communicate with the work machine 200 connected to the power supply device 76 and acquire a signal indicating the status of the work machine 200 from the work machine 200. The control device 4a may also provide the work machine 200 with a signal that controls the operation of the work machine 200. Furthermore, the work machine 200 may generate a signal instructing the operation of the multicopter 10 and transmit it to the control device 4a. Such communication between the control device 4a and the work machine 200 can be performed via wired or wireless communication.

[0035] FIG. 2B is a block diagram showing an example of the basic configuration of a series hybrid drive multicopter 10. Similar to the battery-powered multicopter 10, the series hybrid drive multicopter 10 includes multiple rotors 12, multiple motors 14, multiple ESCs 16, a control device 4a, a sensor group 4b, and a communication device 4c. The illustrated series hybrid drive multicopter 10 further includes an internal combustion engine 7a, a fuel tank 7b for storing fuel for the internal combustion engine 7a, a power generation device 8 driven by the internal combustion engine 7a to generate electric power, a power buffer 9 for temporarily storing the electric power generated by the power generation device 8, and a power supply device 76 electrically connected to the power buffer 9. The power buffer 9 is, for example, a battery such as a secondary battery. The electric power generated by the power generation device 8 is supplied to the motor 14 via the power buffer 9 and the ESC 16. The electric power generated by the power generation device 8 may also be supplied to the work machine 200 via the power supply device 76.

[0036] FIG. 2C is a block diagram showing an example of the basic configuration of a parallel hybrid drive multicopter 10. Similar to the series hybrid drive multicopter 10, the parallel hybrid drive multicopter 10 includes multiple rotors 12, multiple motors 14 that respectively drive the multiple rotors 12, multiple ESCs 16, a control device 4a, a sensor group 4b, a communication device 4c, an internal combustion engine 7a, a fuel tank 7b, a power generator 8, a power buffer 9, and a power supply device 76. The parallel hybrid drive multicopter 10 further includes a drive train 27 that transmits the driving force of the internal combustion engine 7a, and a rotor 22 that rotates by receiving the driving force of the internal combustion engine 7a from the drive train 27. One of the rotor 12 and the rotor 22 may be referred to as the “first rotor” and the other as the “second rotor” to distinguish them from each other. The number of rotors 22 connected to the drive train 27 and rotating may be one or more.

[0037] In the parallel hybrid drive multicopter 10, the internal combustion engine 7a not only drives the power generation device 8 to generate electricity, but also mechanically transmits energy to the rotor 22 to rotate the rotor 22. On the other hand, in the series hybrid drive multicopter 10, all of the rotors 12 are rotated by the electric power generated by the power generation device 8. For this reason, in the series hybrid drive multicopter 10, if the power generation device 8 is, for example, a fuel cell, the internal combustion engine 7a is not an essential component.

[0038] Below, an example of the configuration and operation of an unmanned aerial vehicle according to an embodiment of the present disclosure will be described, taking a multicopter that uses parallel hybrid drive as an example.

[0039] <Basic Configuration> Fig. 3A is a schematic top view of the multicopter 100 according to this embodiment, and Fig. 3B is a side view thereof. Fig. 3B illustrates a work implement 200 coupled to the multicopter 100. In addition to or instead of the work implement 200, luggage, agricultural materials, other machinery, or containers, cases, or packages capable of accommodating these items may be coupled to the multicopter 100. Hereinafter, the weight of the work implement 200 and the work implement itself may be referred to as the "payload." The "coupling" between the multicopter 100 and the work implement 200 or the like may be performed using various tools or devices.

[0040] The multicopter 100 shown in FIG. 3A includes eight sub-rotors 12 and two main rotors 22. Each sub-rotor 12 is composed of four sets of propellers 12a and 12b that rotate coaxially and in opposite directions. Each of the propellers 12a and 12b has two blades. The propellers 12a and 12b are rotated by a motor 14. The four sets of propellers 12a and 12b that rotate coaxially and in opposite directions are located at the vertices of a square. The main rotor 22 is composed of two propellers 22a that rotate in opposite directions at different positions. Each propeller 22a has four blades. The eight propellers 12a and 12b of the sub-rotor 12 have the same pitch angle and diameter. The two propellers 22a of the main rotor 22 also have the same pitch angle and diameter. The diameter of the propeller 22a is 1.2 times or more, for example, 1.4 times or more and 2.0 times or less, the diameter of the propellers 12a and 12b.

[0041] The multicopter 100 includes an airframe 110 having four arms 110A for the sub-rotors 12 and two arms 110B for the main rotors 22. The airframe 110 supports an airframe body 120 including various electronic and mechanical components, which will be described later.

[0042] In the example of FIG. 3B , the airframe main body 120 has a power supply device 76 and an actuator 78 used for coupling to the work machine 200, etc. The power supply device 76 is a device that supplies power generated within the airframe main body 120 to the work machine 200. The actuator 78 is a device such as an electric motor that performs an operation to couple the work machine 200 to the airframe main body 120 of the multicopter 100. In the example of FIG. 3B , the actuator 78 drives a mechanism that winds up a cable connecting the airframe main body 120 and the work machine 200. This cable may include a power supply line for supplying power from the multicopter 100 to the work machine 200, and a communication line for communication between the multicopter 100 and the work machine 200.

[0043] <System Configuration> FIG. 4 is a block diagram showing an example of a system configuration of the multicopter 100 according to this embodiment.

[0044] In the illustrated example, the airframe 120 of the multicopter 100 has a control device 30 including a flight controller 32, a sensor group 72, and a communication device 74. These are basically the same as the control device 4a, the sensor group 4b, and the communication device 4c of the airframe 4 of the multicopter 10 described with reference to FIG. 1A.

[0045] The multicopter 100 of this embodiment includes eight sub-rotors 12, eight motors 14 for rotating the eight sub-rotors 12, and eight ESCs for controlling the eight motors 14. Each ESC 16 receives a signal (motor control signal) for controlling the motor 14 from the control device 30 via a wiring 82. The motor control signal is, for example, a PWM (Pulse With Modulation) signal. When the motor control signal is a PWM signal, the duty cycle of the PWM signal can indicate an analog value of the motor rotation speed. Each ESC 16 controls the rotation speed of the motor 14 connected to that ESC 16 based on the motor control signal from the control device 30. While FIG. 4 illustrates one set of "sub-rotors 12, motors 14, and ESCs 16" for simplicity, the multicopter 100 of this embodiment includes eight sets of "sub-rotors 12, motors 14, and ESCs 16." The number of these sets is not limited to eight.

[0046] The control device 30 is connected to each of the ESCs 16 via electrically independent wiring 82, and can individually control each of the eight ESCs 16. As described above, the sub-rotors 12 are used not only to generate lift but also for attitude control. Attitude control is achieved by the flight controller 32 of the control device 30 obtaining measured or estimated values ​​indicating the attitude of the airframe 120 from the sensors 72, determining the current attitude of the airframe 120, and controlling the rotational speed of each of the motors 14 in accordance with the difference from the target attitude.

[0047] The aircraft body 120 includes a main rotor drive unit 24 that drives the main rotor 22 and a main rotor control unit 26 that controls the main rotor drive unit 24. In this embodiment, the main rotor drive unit 24 is an internal combustion engine. Therefore, the main rotor control unit 26 includes an engine control unit (ECU). The main rotor control unit 26 may also be referred to as a "controller" that controls the internal combustion engine. The main rotor control unit 26 acquires sensor data, such as the accelerator position, intake air temperature, engine speed, and temperatures of various components, of the main rotor drive unit 24, which is an internal combustion engine, and controls the internal combustion engine. The main rotor control unit 26 is connected to the control device 30 via wiring 82, such as a CAN (Controller Area Network) bus. The main rotor control unit 26 is configured to output an engine control signal based on a signal transmitted from the control device 30. The engine control signal includes, for example, a throttle position. A digital-to-analog converter (DAC) and / or a voltage converter may be connected between the control device 30 and the main rotor control unit 26. A mechanical device such as a reducer may be provided between the main rotor drive section 24 and the main rotor 22.

[0048] The main rotor drive unit 24 is preferably an internal combustion engine with little vibration. In this embodiment, the main rotor drive unit 24 is, for example, an opposed-piston engine. Opposed-piston engines are disclosed, for example, in Japanese Patent No. 5508604. The entire contents of Japanese Patent No. 5508604 are incorporated herein by reference.

[0049] The main rotor drive unit 24, which is an internal combustion engine, can generate electricity by driving a power generator 42, such as an alternator. In this embodiment, the power generator 42 has the structure of an AC synchronous motor having a rotor and a stator. Therefore, when the main rotor drive unit 24 is started, the power generator 42 can also function as a "starter" by rotating the rotor when current is applied. The power generator 42 rectifies the AC generated by power generation and converts it into DC. The power generator 42 generates DC power required to drive the motor 14 and supplies it to each ESC 16 via wiring 80. The power generator 42 is configured to output a DC voltage of, for example, 250 V or higher. Note that the wiring 80 is a power wiring, and the wiring 82 is a signal wiring. Each of the wirings 80 and 82 includes multiple conductors.

[0050] The power generation device 42 is connected to a power management device 44. The power management device 44 is connected to the control device 30 and a battery management device 54, which will be described later. The power management device 44 can control the amount of power generated by the power generation device 42 based on signals from the control device 30 or the battery management device 54. This amount of power generation can be variably controlled by the power management device 44 in accordance with the power required by the motor 14 and the battery 52, even when the engine speed of the main rotor drive unit 24, which is an internal combustion engine, is constant.

[0051] The aircraft body 120 further includes a battery 52 in the form of a plurality of cells, for example, lithium ion secondary batteries, connected in series or in parallel, and a battery management device 54 that controls the charging and discharging of the battery 52 .

[0052] The battery 52 receives DC power from the power generation device 42 via a power switch 56 and can be charged by the power. The operation of the power switch 56 can be controlled by a battery management unit 54 and the control device 30. The battery management unit 54 is a device that measures or estimates parameter values ​​that define the state of the battery 52, such as the current flowing through the battery 52, cell voltage, cell balance, state of charge (SOC), state of health (SOH), and temperature.

[0053] The battery management unit 54 can control the power switch 56 depending on the state of the battery 52. ​​For example, when the battery 52 is in a state requiring charging, the battery management unit 54 electrically connects the power generation device 42 and the battery 52 using the power switch 56, and supplies power from the power generation device 42 to the battery 52 to perform a charging operation. At this time, the battery management unit 54 controls the power management unit 44 to increase the amount of power generated by the power generation device 42 so that the power supplied to the ESC 16 does not drop below a desired level. On the other hand, when the battery 52 is in a state requiring no charging, the battery management unit 54 disconnects the electrical connection between the power generation device 42 and the battery 52 using the power switch 56, and stops charging the battery 52.

[0054] In this embodiment, the battery 52 has a storage capacity that allows the aircraft to continue to fly to a location where landing is possible and land there, by continuing to generate lift and control attitude using the sub-rotor 12, even if power generation by the power generation device 42 stops for some reason and lift from the main rotor 22 is lost. In other words, when the multicopter 100 of this embodiment is flying normally, the power required to drive the sub-rotor 12 can be supplied to the ESC 16 from the power generation device 42, rather than from the battery 52. ​​Therefore, even if the payload and flight time are increased, there is little need to increase the storage capacity of the battery 52 accordingly.

[0055] The power stored in the battery 52 can be output as a DC voltage of, for example, 250 V or higher. However, this DC voltage decreases as the charging rate decreases. Therefore, when the charging rate falls below a predetermined level, the battery management device 54 operates to supply part of the DC power from the power generation device 42 to the battery 52 to charge the battery 52.

[0056] The battery 52 is connected to a power circuit board 60. The power circuit board 60 has a function of stepping down the voltage output from the battery 52 to, for example, 24 V, 12 V, or 5 V. The DC voltage output from the battery 52 is converted to a desired voltage by the power circuit board 60 and then supplied to other electronic components.

[0057] 4 , the power supply device 76 is electrically connected to the power generation device 42 or the battery 52 by a power switch 56. The power supply device 76 in this example is configured to supply power generated within the machine body 120 to an external machine or device such as a work machine 200.

[0058] The airframe main body 120 may have a configuration not shown in FIG. 4 . For example, the airframe main body 120 may include a fuel tank that stores fuel necessary for the operation of the main rotor drive unit 24, a water-cooling or air-cooling device for cooling the main rotor drive unit 24, and electrical components such as lighting devices and an electric pump. The electrical components can be operated by power that has been stepped down to a predetermined voltage by the power circuit board 60. A battery for the electrical components (auxiliary battery) may be provided and configured to supply power to the electrical components. Such an auxiliary battery may be charged by the battery 52 or the power generation device 42.

[0059] In this embodiment, the motor 14 functions as a plurality of "attitude control devices" that respectively drive a plurality of first rotors (sub-rotors) 12. In addition, the main rotor drive unit 24, which is an internal combustion engine, functions as a "main thrust generating device" that drives the second rotor (main rotor) 22.

[0060] In this embodiment, the control device 30 is capable of changing the ratio (power ratio) between the first drive power output from the multiple motors 14 and the second drive power output from the main rotor drive unit 24.

[0061] Generally, the responsiveness of the motor 14 is superior to that of an internal combustion engine. If the time from when a torque command signal is input until the torque required to rotate the rotors 12, 22 reaches a target torque value is called the "response time," the response time of the motor is, for example, approximately 1 / 100 of the response time of the internal combustion engine. Therefore, to control the attitude of the multicopter 100, it is desirable to detect the difference between the current and target attitude angles of the multicopter 100 and control the rotational speed of each of the multiple sub-rotors 12 with a high response speed so as to reduce this difference. An increase in rotor rotational speed results in an increase in thrust. By adjusting the thrust of each of the multiple sub-rotors 12, the attitude of the multicopter 100 can be controlled quickly and with high precision.

[0062] On the other hand, an internal combustion engine can efficiently generate large thrust. The sub-rotor 12 is rotated using electric power generated by the power of the main rotor drive unit 24, which is an internal combustion engine. However, energy loss occurs when converting mechanical energy into electrical energy. Therefore, from the viewpoint of improving energy consumption efficiency, it is preferable that the main rotor drive unit 24 be used to rotate the main rotor 22 and generate the main thrust. Furthermore, to increase the thrust of the main rotor 22, it is preferable that the diameter of the main rotor 22 be larger than the diameter of each of the multiple first rotors 12.

[0063] The aircraft main body 120 illustrated in Figure 4 further includes a decompression control device 91, a water pump 92, and a fuel pump 93. The decompression control device 91 is a device for controlling the decompression operation provided in the main rotor drive unit 24, which is an internal combustion engine. An example of the decompression control device 91 is a solenoid. A solenoid is a type of motor.

[0064] The aircraft body 120 is provided with a water-cooling system for cooling the main rotor drive unit 24. The water-cooling system includes, for example, a water pump 92, a radiator, a radiator fan, and a thermostat. The water pump 92 is an electrical component for controlling the circulation of coolant (cooling water) supplied to the main rotor drive unit 24. The fuel pump 93 is an electrical component for supplying fuel from a fuel tank to the main rotor drive unit 24.

[0065] 4 , the decompression control device 91, the water pump 92, and the fuel pump 93 are each connected to the power circuit board 60 via wiring 80 and receive power from the power circuit board 60. For example, a relay that operates under the control of the main rotor control unit 26 may be provided on the wiring 80 that connects the decompression control device 91, the water pump 92, and the fuel pump 93, respectively. The desired power that has been stepped down by the power circuit board 60 can be supplied to these electrical components from the power circuit board 60 via the relay. The power that has been stepped down by the power circuit board 60 is also supplied to the control device 30 and the actuator 78 via the wiring 80.

[0066] 4 further includes a relay 94 connected between the main rotor control unit 26 and the power circuit board 60. Desired power is supplied from the power circuit board 60 to the main rotor control unit 26 via the relay 94. The relay 94 is electrically connected to the control device 30 via wiring 82 and operates under the control of the control device 30. By turning off the relay 94, the electrical connection between the main rotor control unit 26 and the power circuit board 60 can be cut off, and the power supply to the main rotor control unit 26 can be instantly stopped.

[0067] FIG. 5 is a block diagram showing an example of a power system in which power is supplied from a battery 52 to each electrical component.

[0068] As described above, the power circuit board 60 has the function of stepping down the voltage output from the battery 52 to, for example, 24 V, 12 V, or 5 V. In this embodiment, the battery voltage of the battery 52 is 288 V. In other words, the DC voltage output from the battery 52 is 288 V. However, this voltage value is an example, and the DC voltage may be, for example, 250 V or higher. The power circuit board 60 may further include a first power supply circuit that steps down a first DC voltage based on the battery voltage of the battery 52 to a second DC voltage, and a second power supply circuit that steps down the battery voltage of the battery 52 to the first DC voltage. Each of the first and second power supply circuits may be, for example, a DC-DC converter. The first DC voltage may be, for example, 20 V or higher and 30 V or lower, and the second DC voltage may be, for example, 10 V or higher and 20 V or lower. The ratio of the first DC voltage to the second DC voltage may be, for example, 1.5 or higher and 3 or lower.

[0069] In this embodiment, the battery voltage is 288 V, the first DC voltage is 24 V, and the second DC voltage is 12 V. In the example shown in FIG. 5 , the power circuit board 60 includes a DC-DC converter 61 that steps down the 288 V battery voltage to a first DC voltage of 24 V, and a DC-DC converter 62 that steps down the 24 V first DC voltage to a second DC voltage of 12 V. The DC-DC converter 61 corresponds to the second power supply circuit, and the DC-DC converter 62 corresponds to the first power supply circuit. The power circuit board 60 may further include a DC-DC converter that steps down the 24 V first DC voltage or the 12 V second DC voltage to a 5 V DC voltage. The 5 V DC voltage is used as a power supply voltage for electronic components such as the flight controller 32 included in the control device 30.

[0070] A second DC voltage of 12 V is supplied to a plurality of electrical components from DC-DC converter 62 of power circuit board 60. In the example shown in Fig. 5, the plurality of electrical components include main rotor control unit 26 (ECU), a decompression control device 91, a water pump 92, a fuel pump 93, and a relay 94. The second DC voltage of 12 V is supplied from DC-DC converter 62 to these electrical components connected to DC-DC converter 62.

[0071] The multiple electrical components connected to the DC-DC converter 62 include a first group of electrical components that consume less power than a threshold power when activated, and a second group of electrical components that consume more power than a threshold power when activated. "Activation" refers to supplying power to the electrical components to initiate a desired operation. In the example shown in FIG. 5 , the first group of electrical components includes a relay 94 and a main rotor control unit 26. The second group of electrical components includes a decompression control device 91, a water pump 92, and a fuel pump 93. The second group of electrical components further includes a first electrical component that consumes a first power when activated, and a second electrical component that consumes a second power when activated. The first power may be the same as or different from the second power. The second group of electrical components further includes a third electrical component that consumes a third power when activated that is less than the first power and less than the second power. In the example shown in FIG. 5, the first electrical component is a fuel pump 93 , the second electrical component is a decompression control device 91 , and the third electrical component is a water pump 92 .

[0072] The electric motors included in the second electrical component group, such as the decompression control device (solenoid) 91, the water pump 92, and the fuel pump 93, require a large amount of power when they are started. In other words, these electrical components consume a large amount of power when they are started. This is because an inrush current temporarily flows through these electrical components when they are started. When comparing the power consumed when they are started, the power consumed by the decompression control device 91, the water pump 92, and the fuel pump 93 is greater than the power consumed by either the relay 94 or the main rotor control unit 26.

[0073] In this embodiment, the power consumption during startup is, for example, 2 W, 24 W, 144 W, 72 W, and 180 W for the relay 94, main rotor control unit 26, decompression control device 91, water pump 92, and fuel pump 93, in that order. The fuel pump 93, an example of a first electrical component, consumes a first power of 180 W during startup. The decompression control device 91, an example of a second electrical component, consumes a second power of 144 W during startup. The water pump 92, an example of a third electrical component, consumes a third power of 72 W during startup. In this way, the relationship of first power > second power > third power holds. The threshold power described above is set, for example, in the range of 30 W to 70 W.

[0074] The steady-state power consumption of the relay 94, main rotor control unit 26, decompression control device (solenoid) 91, water pump 92, and fuel pump 93, in that order, is, for example, 2 W, 24 W, 0 W, 72 W, and 72 W. The solenoid generates an attractive force while its internal coil is energized to operate the decompression, and consumes power at the same time. Therefore, the solenoid does not consume power during steady-state operation.

[0075] The rated power of the first power supply circuit is, for example, 200 W or more and 400 W or less. The rated power of the second power supply circuit is, for example, 500 W or more and 700 W or less. The ratio of the rated power of the second power supply circuit to the rated power of the first power supply circuit is preferably, for example, 1.5 or more and 3.5 or less. In this embodiment, the rated power of the DC-DC converter 62, which is the first power supply circuit, is 300 W, and the rated power of the DC-DC converter 61, which is the second power supply circuit, is 600 W. Consider the case where all of the first and second electrical component groups connected to the DC-DC converter 62, which is the first power supply circuit, are started simultaneously. In this case, the total power consumption during startup is 422 W, which exceeds the rated power (300 W) of the DC-DC converter 62. Therefore, to start all of the first and second electrical component groups simultaneously, a larger power supply circuit with a rated power of 422 W or more is required. To solve this problem, in this embodiment, the timing at which the first electrical component included in the second electrical component group starts after the main rotor drive unit 24 starts is made different from the timing at which the second electrical component included in the second electrical component group starts.

[0076] In this embodiment, the control device 30 controls the timing of startup of the first and second electrical component groups. In other words, the control device 30 controls the startup sequence of the first and second electrical component groups. Specifically, after the main rotor drive unit 24 starts (e.g., after the ignition switch is turned on), the control device 30 starts the first electrical component group and the first electrical component, and then starts the second electrical component. After starting some of the electrical components included in the second electrical component group (e.g., the second electrical component), the control device 30 further starts other electrical components included in the second electrical component group (e.g., the third electrical component).

[0077] FIG. 8 is a graph showing the change over time in the power consumption of the DC-DC converter 62, the first electrical component, and the second electrical component. In FIG. 8, the top graph (A) shows the power consumption of the DC-DC converter 62, the middle graph (B) shows the power consumption of the first electrical component, and the bottom graph (C) shows the power consumption of the second electrical component. The vertical axis of each graph represents power consumption, and the horizontal axis represents time. As shown in FIG. 8, the control device 30 starts up the first group of electrical components and the first electrical component, and then starts up some or all of the second group of electrical components after the power consumption of the first electrical component decreases. This startup sequence makes it possible to keep the total power consumption of the multiple electrical components connected to the DC-DC converter 62 within the rated power of the DC-DC converter 62.

[0078] 6 is a flowchart illustrating the startup sequence of the first and second electrical component groups. The control device 30 first starts the relay 94 and main rotor control unit 26 included in the first electrical component group, and the fuel pump 93, which is the first electrical component included in the second electrical component group (step S10). The total power consumption of the relay 94, main rotor control unit 26, and fuel pump 93 during startup is 206 W, which is less than the rated power (300 W) of the DC-DC converter 61.

[0079] Next, after activating the relay 94, the main rotor control unit 26, and the fuel pump 93, the control device 30 activates the decompression control device 91, which is a second electrical component included in the second electrical component group, for example, two seconds later (step S20). During the decompression control device 91 activation phase, the activated relay 94, main rotor control unit 26, and fuel pump 93 each consume power in a steady state. Therefore, the power consumption of the electrical components corresponds to the sum of the steady-state power consumption of the relay 94, main rotor control unit 26, and fuel pump 93 and the power consumption of the decompression control device 91 during activation. This total power consumption is 242 W, which is less than the rated power (300 W) of the DC-DC converter 61. Thus, in this embodiment, during the three phases of steps S10 to S30, the decompression control device 91, which does not consume power in a steady state, is activated in the phase between steps S10 and S30 (step S20).

[0080] Next, after starting the decompression control device 91, for example, one second later, the control device 30 starts the water pump 92, which is a third electrical component included in the second electrical component group (step S30). During the phase of starting the water pump 92, the relay 94, main rotor control unit 26, and fuel pump 93, which are already activated, each consume electric power in a steady state. As described above, the decompression control device 91 does not consume electric power in a steady state. Therefore, the power consumption of the electrical components corresponds to the sum of the power consumption of the relay 94, main rotor control unit 26, and fuel pump 93 in a steady state and the power consumption of the water pump 92 when it is started. This total power consumption is 170 W, which is less than the rated power (300 W) of the DC-DC converter 61.

[0081] After the water pump 92 is started, all of the electrical components included in the first and second electrical component groups consume electric power in a steady state. The total electric power consumed by the first and second electrical component groups in a steady state illustrated in FIG. 5 is 170 W, which is less than the rated power (300 W) of the DC-DC converter 61.

[0082] According to the startup sequence of this embodiment, when starting up a plurality of electrical components connected to the first power supply circuit, the startup timings of the first to third electrical components included in the second electrical component group are adjusted so that power peaks caused by inrush currents temporarily flowing through the first to third electrical components do not overlap in time. Specifically, the startup timing of the second electrical component is delayed from the startup timing of the first electrical component, and the startup timing of the third electrical component is delayed from the startup timing of the second electrical component, thereby shifting the respective power peaks. This allows the power consumption to be less than the rated power of the first power supply circuit, even if the total power consumed by the plurality of electrical components during startup exceeds the rated power. As a result, it becomes possible to use a power supply circuit (DC-DC converter) with a lower rated power. This can contribute to reducing the size, cost, and weight of the power supply circuit. Lightweight multicopters are desirable in order to increase flight time. In particular, reducing the weight of the power supply circuit can contribute to reducing the weight of the multicopters.

[0083] 7 is a block diagram showing an example of the hardware configuration of the control device 30. The control device 30 includes a processor 34, a ROM (Read Only Memory) 35, a RAM (Random Access Memory) 36, a storage device 37, and a communication I / F 38. These components are connected to each other via a bus 39.

[0084] The processor 34 is one or more semiconductor integrated circuits, and is also referred to as a central processing unit (CPU) or a microprocessor. The processor 34 sequentially executes computer programs stored in the ROM 35 to perform the above-described processing. The term "processor 34" is broadly interpreted as including a field programmable gate array (FPGA) equipped with a CPU, a graphic processor unit (GPU), an application specific integrated circuit (ASIC), or an application specific standard product (ASSP).

[0085] The ROM 35 is, for example, a writable memory (e.g., a PROM), a rewritable memory (e.g., a flash memory), or a read-only memory. The ROM 35 stores a program that controls the operation of the processor. The ROM 35 does not have to be a single recording medium, but may be a collection of multiple recording media. Some of the collection of multiple recording media may be removable memories.

[0086] The RAM 36 provides a working area for temporarily loading the programs stored in the ROM 35 at boot time. The RAM 36 does not have to be a single recording medium, but can be a collection of multiple recording media.

[0087] The communication I / F 38 is an interface for communication between the control device 30 and other electronic components or electronic control units (ECUs). For example, the communication I / F 38 can perform wired communication in accordance with various protocols. The communication I / F 38 may also perform wireless communication in accordance with the Bluetooth (registered trademark) standard and / or the Wi-Fi (registered trademark) standard. Both standards include wireless communication standards that use frequencies in the 2.4 GHz band.

[0088] The storage device 37 may be, for example, a semiconductor memory, a magnetic storage device, an optical storage device, or a combination thereof. The storage device 37 may store, for example, map data useful for the autonomous flight of the multicopter 10 and various sensor data acquired by the multicopter 10 during flight.

[0089] As described above, the control device 4a may include, for example, a flight control device such as a flight controller and a host computer (companion computer). The companion computer may execute the processes shown in FIG. 6 and issue commands related to a startup sequence based on the results of those processes to the control device 30. Some or all of the functions of the electrical components, such as the control device 30, mounted on the multicopter 100 may be implemented by one or more servers (computers) or terminal devices (including portable and fixed types) connected to the communication device 4c of the multicopter 100 via a communication network.

[0090] A system providing various functions in the embodiments can also be retrofitted to a multicopter that does not have those functions. Such a system can be manufactured and sold independently of the multicopter. A computer program used in such a system can also be manufactured and sold independently of the multicopter. The computer program can be provided, for example, by being stored in a computer-readable non-transitory storage medium. The computer program can also be provided by downloading via a telecommunications line (e.g., the Internet).

[0091] This specification discloses the solutions described in the following items.

[0092] [Item 1] An unmanned aerial vehicle having multiple rotors, comprising: an internal combustion engine; a power generation device driven by the internal combustion engine to generate electric power; a battery that stores the electric power; and multiple electrical components to which electric power is supplied from the battery, wherein the multiple electrical components include a first group of electrical components that consume less electric power than a threshold power at startup, and a second group of electrical components that consume more electric power than the threshold power at startup, and the second group of electrical components includes a first electrical component that consumes a first power at startup, and a second electrical component that consumes a second power at startup, and the timing at which the first electrical components start up after the internal combustion engine starts is different from the timing at which the second electrical components start up.

[0093] [Item 2] The unmanned aerial vehicle according to Item 1, wherein the second electrical component is activated after the first electrical component group and the first electrical component are activated after the internal combustion engine is started.

[0094] [Item 3] The unmanned aerial vehicle according to item 1 or 2, wherein the second electrical component does not consume power in a steady state.

[0095] [Item 4] An unmanned aerial vehicle described in any one of items 1 to 3, comprising a plurality of electric motors that respectively drive a plurality of first rotors included in the plurality of rotors, and the internal combustion engine drives at least one second rotor included in the plurality of rotors.

[0096] [Item 5] The unmanned aerial vehicle according to Item 4, wherein the second electrical component is a decompression control device that controls a decompression operation provided in the internal combustion engine.

[0097] [Item 6] The unmanned aerial vehicle according to Item 5, further comprising: a controller that controls the internal combustion engine; and the first group of electrical components includes the controller.

[0098] [Item 7] The unmanned aerial vehicle described in Item 6, wherein the second group of electrical components further includes a third electrical component that consumes a third power at startup that is smaller than the first power and smaller than the second power, and the second electrical component starts up after the internal combustion engine starts and the first group of electrical components and the first electrical component start up, and the third electrical component starts up after the second electrical component starts up.

[0099] [Item 8] An unmanned aerial vehicle as described in Item 7, comprising: a water pump that controls the circulation of coolant supplied to the internal combustion engine; and a fuel pump that supplies fuel to the internal combustion engine, wherein the first electrical component is the fuel pump and the third electrical component is the water pump.

[0100] [Item 9] An unmanned aerial vehicle described in any one of items 1 to 8, comprising a first power supply circuit that steps down a first DC voltage based on the battery voltage of the battery to a second DC voltage.

[0101] [Item 10] The unmanned aerial vehicle according to Item 9, further comprising a second power supply circuit that steps down the battery voltage of the battery to the first DC voltage.

[0102] [Item 11] The unmanned aerial vehicle described in Item 10, wherein the ratio of the rated power of the second power supply circuit to the rated power of the first power supply circuit is 1.5 or more, and the ratio of the first DC voltage to the second DC voltage is 1.5 or more.

[0103] The unmanned aerial vehicle disclosed herein can be widely used not only for aerial photography, surveying, logistics, and pesticide spraying, but also for ground work related to agricultural work, transporting harvested products and agricultural materials, and the like.

[0104] 2: Rotor (propeller), 3: Rotation drive device, 4: Airframe body, 4a: Control device, 4b: Sensor group, 4c: Communication device, 5: Airframe frame, 10: Multicopter, 12: Sub-rotor, 12a: Propeller, 12b: Propeller, 14: Motor, 16: ESC, 22: Main rotor, 52: Battery, 54: Battery management device

Claims

1. 1. An unmanned aerial vehicle having multiple rotors, an internal combustion engine; a power generation device driven by the internal combustion engine to generate electric power; a battery that stores the power; a plurality of electrical components to which power is supplied from the battery; Equipped with the plurality of electrical components include a first group of electrical components that consume less power than a threshold power when activated, and a second group of electrical components that consume more power than the threshold power when activated, the second group of electrical components includes a first electrical component that consumes a first amount of power when activated and a second electrical component that consumes a second amount of power when activated; The unmanned aerial vehicle has a timing at which the first electrical component is activated after the internal combustion engine is started that is different from a timing at which the second electrical component is activated.

2. The unmanned aerial vehicle according to claim 1 , wherein the second electrical component is activated after the first electrical component group and the first electrical component are activated after the internal combustion engine is started.

3. The unmanned aerial vehicle according to claim 1 or 2, wherein the second electrical component does not consume power in a steady state.

4. a plurality of electric motors that respectively drive a plurality of first rotors included in the plurality of rotors; The unmanned aerial vehicle according to claim 1 or 2, wherein the internal combustion engine drives at least one second rotor included in the plurality of rotors.

5. The unmanned aerial vehicle according to claim 4 , wherein the second electrical component is a decompression control device that controls a decompression operation provided in the internal combustion engine.

6. a controller for controlling the internal combustion engine; Equipped with The unmanned aerial vehicle according to claim 5 , wherein the first group of electrical components includes the controller.

7. the second group of electrical components further includes a third electrical component that consumes a third power during startup that is smaller than the first power and smaller than the second power; 7. The unmanned aerial vehicle described in claim 6, wherein the internal combustion engine is started, the first group of electrical components and the first electrical component are activated, and then the second electrical component is activated, and the third electrical component is activated after the second electrical component is activated.

8. a water pump that controls the circulation of coolant supplied to the internal combustion engine; a fuel pump for supplying fuel to the internal combustion engine; Equipped with The unmanned aerial vehicle according to claim 7 , wherein the first electrical component is the fuel pump, and the third electrical component is the water pump.

9. 3. The unmanned aerial vehicle according to claim 1, further comprising a first power supply circuit that steps down a first DC voltage based on the battery voltage of the battery to a second DC voltage.

10. The unmanned aerial vehicle according to claim 9 , further comprising a second power supply circuit that steps down the battery voltage of the battery to the first DC voltage.

11. a ratio of the rated power of the second power supply circuit to the rated power of the first power supply circuit is 1.5 or more; The unmanned aerial vehicle of claim 10 , wherein the ratio of the first DC voltage to the second DC voltage is 1.5 or greater.