Control device for an aircraft

Abstract

A control device of an aircraft is provided. The control device (48) of the aircraft (10) controls the force generated by each VTOL rotor (12) in a state where the rotation of the VTOL rotor (12) is stopped by controlling the pitch of each blade (26) and the rotation stop angle of each VTOL rotor (12) in a state where the lift of the front wing (18) and the rear wing (20) is generated, and thereby the force is applied to the fuselage (16). Accordingly, in a state where the rotation of the rotor is stopped, better control can be provided when the force generated by the rotor is used to control the force applied to the fuselage.

Description

Technical Field

[0001] This invention relates to a control device for an aircraft. Background Technology

[0002] U.S. Patent Application Publication No. 2020 / 0079501 ​​discloses a control device for an aircraft. Unlike wings that generate lift, this aircraft has a rotor that generates lift through rotation. During cruise, the control device positions the rotor blades in an optimal position, thereby minimizing airflow turbulence over the wings during cruise. Summary of the Invention

[0003] Better control is needed when using the forces generated in a rotor that has stopped rotating to control the forces acting on the fuselage.

[0004] The purpose of this invention is to solve the above-mentioned technical problems.

[0005] The present invention relates to a control device for an aircraft having a plurality of rotors and one or more fixed wings, wherein each of the plurality of rotors has one or more blades and generates lift by rotating about a rotation axis; the one or more fixed wings generate lift when the fuselage has airspeed, and the control device is characterized in that, when the fixed wings are generating lift, the force generated by each rotor when the rotation of the rotors is stopped is controlled by controlling the pitch of each of the blades and the rotation stop angle of each rotor, thereby causing the force to act on the fuselage.

[0006] According to the present invention, the force generated in the rotor is used when the rotor has stopped rotating, which provides better control when controlling the forces acting on the fuselage.

[0007] The above-described objectives, features, and advantages should be readily understood from the following description of the embodiments with reference to the accompanying drawings. Attached Figure Description

[0008] Figure 1 This is a schematic diagram of an aircraft.

[0009] Figure 2 It is a diagram showing the structure of an electric power supply system.

[0010] Figure 3 This is a block diagram of the control device. Detailed Implementation

[0011] [First Embodiment] [Aircraft Structure] Figure 1This is a schematic diagram of aircraft 10. Aircraft 10 in this embodiment is an electric vertical takeoff and landing (eVTOL) aircraft. Aircraft 10 in this embodiment has multiple VTOL rotors 12 and multiple cruise rotors 14. Rotation of the VTOL rotors 12 generates lift that moves the fuselage 16 upwards. Rotation of the cruise rotors 14 generates thrust that moves the fuselage 16 forwards.

[0012] Aircraft 10 has a fuselage 16. A cockpit, passenger cabin, etc., are located in the fuselage 16. The pilot sits in the cockpit and controls the aircraft 10. Passengers and other passengers sit in the passenger cabin. Aircraft 10 can also be controlled automatically without a pilot.

[0013] Aircraft 10 has a canard 18 and a rear wing 20 as fixed wings. The canard 18 is positioned forward of the center of gravity G of the fuselage 16. The rear wing 20 is positioned aft of the center of gravity G of the fuselage 16. When the fuselage 16 has an airspeed, the angle of attack of the canard 18 and the rear wing 20 is controlled, thereby generating lift at the canard 18 and the rear wing 20.

[0014] At least one of the canard wing 18 and the aft wing 20 is provided with an aileron (not shown), an elevator, and a rudder. By operating the aileron, a roll moment can be applied about the center of gravity G of the fuselage 16. By operating the elevator, a pitch moment can be applied about the center of gravity G of the fuselage 16. By operating the rudder, a yaw moment can be applied about the center of gravity G of the fuselage 16.

[0015] The aircraft 10 has eight VTOL rotors 12. The eight VTOL rotors 12 are rotor 12FLa, rotor 12FLb, rotor 12RLa, rotor 12RLb, rotor 12FRa, rotor 12FRb, rotor 12RRa, and rotor 12RRb. The VTOL rotors 12 are respectively equivalent to the rotors of the present invention.

[0016] Rotors 12FLa, 12FLb, 12RLa, and 12RLb are mounted on cantilever 22L. Cantilever 22L extends in the fore-and-aft direction. Cantilever 22L is mounted on the fore wing 18 and the rear wing 20. Cantilever 22L is positioned to the left of the center of gravity G. That is, rotors 12FLa, 12FLb, 12RLa, and 12RLb are positioned to the left of the center of gravity G.

[0017] Rotor 12FLa is positioned forward of canard 18. Rotor 12FLb is positioned between canard 18 and center of gravity G in the longitudinal direction of fuselage 16. Rotor 12RLb is positioned between center of gravity G and rear wing 20 in the longitudinal direction of fuselage 16. Rotor 12RLa is positioned rearward of rear wing 20. The distance from center of gravity G to rotor 12FLb is shorter than the distance from center of gravity G to rotor 12FLa. The distance from center of gravity G to rotor 12RLb is shorter than the distance from center of gravity G to rotor 12RLa.

[0018] Rotor 12FLa corresponds to the outer rotor and the first rotor of the present invention. Rotor 12FLb corresponds to the inner rotor and the second rotor of the present invention. Rotor 12RLb corresponds to the inner rotor and the third rotor of the present invention. Rotor 12RLa corresponds to the outer rotor and the fourth rotor of the present invention.

[0019] Rotors 12FRa, 12FRb, 12RRa, and 12RRb are mounted on cantilever 22R. Cantilever 22R extends in the fore-and-aft direction. Cantilever 22R is mounted on the fore wing 18 and the rear wing 20. Cantilever 22R is positioned to the right of the center of gravity G. That is, rotors 12FRa, 12FRb, 12RRa, and 12RRb are positioned to the right of the center of gravity G.

[0020] Rotor 12FRa is positioned forward of canard 18. Rotor 12FRb is positioned between canard 18 and center of gravity G in the longitudinal direction of fuselage 16. Rotor 12RRb is positioned between center of gravity G and rear wing 20 in the longitudinal direction of fuselage 16. Rotor 12RRa is positioned rearward of rear wing 20. The distance from center of gravity G to rotor 12FRb is shorter than the distance from center of gravity G to rotor 12FRa. The distance from center of gravity G to rotor 12RRb is shorter than the distance from center of gravity G to rotor 12RRa.

[0021] Rotor 12FRa corresponds to the outer rotor and the first rotor of the present invention. Rotor 12FRb corresponds to the inner rotor and the second rotor of the present invention. Rotor 12RRb corresponds to the inner rotor and the third rotor of the present invention. Rotor 12RRa corresponds to the outer rotor and the fourth rotor of the present invention.

[0022] exist Figure 1In the original design, cantilever 22L and cantilever 22R extend in a straight line along the longitudinal direction of fuselage 16. However, cantilever 22L and cantilever 22R can also be formed as arcs protruding outwards in the lateral direction of fuselage 16. When cantilever 22L is formed as an arc protruding outwards in the lateral direction of fuselage 16, rotor 12FLb is located to the left (outer) of rotor 12FLa in the lateral direction of fuselage 16. When cantilever 22R is formed as an arc protruding outwards in the lateral direction of fuselage 16, rotor 12FRb is located to the right (outer) of rotor 12FRa in the lateral direction of fuselage 16.

[0023] Each VTOL rotor 12 has a rotation shaft 24. The rotation shaft 24 extends along the vertical direction of the fuselage 16. The rotation shaft 24 may also have an angle (tilt angle) of several degrees relative to the vertical direction of the fuselage 16.

[0024] Each VTOL rotor 12 has three blades 26. The number of blades 26 of the VTOL rotor 12 is not limited to three. The number of blades 26 of the VTOL rotor 12 can be one or more.

[0025] The VTOL rotor 12 rotates around the rotation axis 24, thereby generating lift at the blade 26. The magnitude of the lift generated at the VTOL rotor 12 is controlled by adjusting the rotation speed and the pitch of the blade 26. The VTOL rotor 12 mainly rotates to generate lift during vertical takeoff, transition from vertical takeoff to cruise, transition from cruise to vertical landing, vertical landing, and hovering.

[0026] When viewed from above, rotors 12FLa, 12RLa, 12FRb, and 12RRb rotate to the left. When viewed from above, rotors 12FRa, 12RRa, 12FLb, and 12RLb rotate to the right. Furthermore, the rotation direction of each VTOL rotor 12 is not limited to the directions described above.

[0027] Even when the VTOL rotor 12 is in a stopped state, it still generates lift when the fuselage 16 has an airspeed. The magnitude of the lift generated by the VTOL rotor 12 is controlled by adjusting the stop angle and the pitch of the blades 26. Similarly, the magnitude of the drag generated by the VTOL rotor 12 is controlled by adjusting the stop angle and the pitch of the blades 26. For example, during cruise, the magnitude of lift and drag generated by the VTOL rotor 12 can be controlled even when it is in a stopped state. Furthermore, not limited to cruise, the magnitude of lift and drag generated by the VTOL rotor 12 can also be controlled during transitions from vertical takeoff to cruise, and from cruise to vertical landing, even when it is in a stopped state. Additionally, when the fuselage 16 has an airspeed and the canard 18 and aft 20 generate lift, the magnitude of lift and drag generated by the VTOL rotor 12 can also be controlled even when it is in a stopped state.

[0028] By controlling the magnitude of lift and drag generated by each of the eight VTOL rotors 12, it is possible to generate roll moment, pitch moment and yaw moment around the center of gravity G of the fuselage 16.

[0029] By creating a difference between the sum of the lift forces of the four VTOL rotors 12 located to the left of the center of gravity G and the sum of the lift forces of the four VTOL rotors 12 located to the right of the center of gravity G, a rolling torque can be applied around the center of gravity G of the fuselage 16.

[0030] The four VTOL rotors 12 located to the left of the center of gravity G are rotors 12FLa, 12FLb, 12RLa, and 12RLb. The four VTOL rotors 12 located to the right of the center of gravity G are rotors 12FRa, 12FRb, 12RRa, and 12RRb.

[0031] By creating a difference between the sum of the lift forces of the four VTOL rotors 12 positioned in front of the center of gravity G and the sum of the lift forces of the four VTOL rotors 12 positioned behind the center of gravity G, a pitching moment can be applied around the center of gravity G of the main body 16.

[0032] The four VTOL rotors 12 positioned in front of the center of gravity G are designated as rotor 12FLa, rotor 12FLb, rotor 12FRa, and rotor 12FRb. The four VTOL rotors 12 positioned behind the center of gravity G are designated as rotor 12RLa, rotor 12RLb, rotor 12RRa, and rotor 12RRb.

[0033] By creating a difference between the sum of the drag values ​​of the four VTOL rotors 12 located to the left of the center of gravity G and the sum of the drag values ​​of the four VTOL rotors 12 located to the right of the center of gravity G, a yawing moment can be applied around the center of gravity G of the fuselage 16.

[0034] The aircraft 10 has two cruise rotors 14. The two cruise rotors 14 are rotor 14L and rotor 14R.

[0035] Rotors 14L and 14R are mounted at the rear of fuselage 16. Rotor 14L is positioned to the left of centerline A of fuselage 16. Rotor 14R is positioned to the right of centerline A of fuselage 16.

[0036] The rotation axis (not shown) of cruise rotor 14 extends along the fore-and-aft direction of fuselage 16. The rotation axis of cruise rotor 14 may also have an angle (tilt angle) of several degrees relative to the fore-and-aft direction. Cruise rotor 14 has one or more blades (not shown).

[0037] Cruise rotor 14 rotates around its rotation axis, thereby generating thrust in its blades. The magnitude of the thrust is controlled by adjusting the rotational speed and the pitch of the blades. Cruise rotor 14 primarily rotates to generate thrust during transitions from vertical takeoff to cruise, during cruise, and during transitions from cruise to vertical landing.

[0038] [Structure of the power supply system] Figure 2 This is a diagram showing the structure of the electrical power supply system 28.

[0039] Each VTOL rotor 12 is equipped with a VTOL motor 30, an inverter 32, and a pitch drive mechanism 34 as a drive mechanism.

[0040] The VTOL motor 30 is PWM controlled by the inverter 32. This controls the rotational speed of the VTOL rotor 12. The pitch drive mechanism 34 changes the pitch of the blades 26 of the VTOL rotor 12.

[0041] Additionally, the rotation stop angle of the VTOL rotor 12 is controlled by PWM control of the VTOL motor 30. The VTOL rotor 12 maintains a predetermined rotation angle in the rotation stop state. The rotation stop angle represents the predetermined rotation angle that is maintained. In order to keep the VTOL rotor 12 stopped at the rotation stop angle, the inverter 32 controls the VTOL motor 30 to output a balanced torque relative to the external force.

[0042] The high-voltage electrical power of several hundred volts output from the power supply unit 36 ​​is supplied to the VTOL motor 30. The electrical power output from the power supply unit 36 ​​is stepped down to a lower voltage of tens of volts in the DC / DC converter 38 and then supplied to the pitch drive mechanism 34.

[0043] Each cruise rotor 14 is equipped with a cruise motor 40, an inverter 42, and a pitch drive mechanism 44 as its drive mechanism.

[0044] Cruise motor 40 is PWM controlled by inverter 42. This controls the rotational speed of cruise rotor 14. Pitch drive mechanism 44 changes the pitch of the blades of cruise rotor 14.

[0045] The higher voltage of several hundred volts of electrical power output from the power supply unit 36 ​​is supplied to the cruise motor 40. The electrical power output from the power supply unit 36 ​​is stepped down to a lower voltage of tens of volts in the DC / DC converter 46 and then supplied to the pitch drive mechanism 44.

[0046] The power supply unit 36 ​​includes an engine (not shown), a generator, a power control unit, and a battery. The engine drives the generator to generate electricity. The power control unit converts the alternating current (AC) power generated by the generator into direct current (DC) power and outputs it. The battery stores a portion of the electrical power output from the power control unit. When the electrical power generated by the generator is insufficient to meet the power consumption of the VTOL motor 30, cruise motor 40, etc., the battery outputs electrical power.

[0047] The power supply unit 36 ​​may have an engine, a generator, and a power control unit, or it may not have a battery. Alternatively, the power supply unit 36 ​​may have a battery but not an engine, generator, and power control unit.

[0048] Inverter 32, pitch drive mechanism 34, inverter 42 and pitch drive mechanism 44 are controlled by control device 48.

[0049] [Structure of the control device] Figure 3 This is a block diagram of the control device 48. The control device 48 has an arithmetic unit 50 and a storage unit 52.

[0050] The arithmetic unit 50 is, for example, a CPU (Central Processing Unit) or a GPU (Graphics Processing Unit). The arithmetic unit 50 includes a lift command value calculation unit 54, a roll moment command value calculation unit 56, a pitch moment command value calculation unit 58, a yaw moment command value calculation unit 60, a mode selection unit 62, a thrust command value calculation unit 64, a distribution command value calculation unit 66, a control surface control command value generation unit 68, a control surface control unit 70, a VTOL rotor command value generation unit 72, a VTOL motor control unit 74, a pitch control unit 76, a cruise rotor command value generation unit 78, a cruise motor control unit 80, and a pitch control unit 82.

[0051] The lift command value calculation unit 54, roll moment command value calculation unit 56, pitch moment command value calculation unit 58, yaw moment command value calculation unit 60, mode selection unit 62, thrust command value calculation unit 64, distribution command value calculation unit 66, control surface control command value generation unit 68, control surface control unit 70, VTOL rotor command value generation unit 72, VTOL motor control unit 74, pitch control unit 76, cruise rotor command value generation unit 78, cruise motor control unit 80, and pitch control unit 82 are implemented by the calculation unit 50 executing the program stored in the storage unit 52.

[0052] The lift command value calculation unit 54, roll torque command value calculation unit 56, pitch torque command value calculation unit 58, yaw torque command value calculation unit 60, mode selection unit 62, thrust command value calculation unit 64, distribution command value calculation unit 66, control surface control command value generation unit 68, control surface control unit 70, VTOL rotor command value generation unit 72, VTOL motor control unit 74, pitch control unit 76, cruise rotor command value generation unit 78, cruise motor control unit 80, and pitch control unit 82 can also be implemented using integrated circuits such as ASIC (Application Specific Integrated Circuit) and FPGA (Field-Programmable Gate Array).

[0053] At least a portion of the lift command value calculation unit 54, roll moment command value calculation unit 56, pitch moment command value calculation unit 58, yaw moment command value calculation unit 60, mode selection unit 62, thrust command value calculation unit 64, distribution command value calculation unit 66, control surface control command value generation unit 68, control surface control unit 70, VTOL rotor command value generation unit 72, VTOL motor control unit 74, pitch control unit 76, cruise rotor command value generation unit 78, cruise motor control unit 80, and pitch control unit 82 can also be implemented by electronic circuits containing discrete components.

[0054] The storage unit 52 comprises volatile memory (not shown) and non-volatile memory (not shown), which are computer-readable storage media. The volatile memory is, for example, RAM (Random Access Memory). The non-volatile memory is, for example, ROM (Read Only Memory), flash memory, etc. Data is stored, for example, in the volatile memory. Programs, tables, maps, etc., are stored, for example, in the non-volatile memory. At least a portion of the storage unit 52 may also be disposed in the aforementioned processor, integrated circuit, etc.

[0055] The lift command calculation unit 54 calculates the lift command value. For example, the lift command value is determined based on the amount of operation performed by the pilot on the operation input unit. The operation input unit may be, for example, a joystick, pedal, or handle. The operation amount of the operation input unit and the lift command value may not be a one-to-one correspondence. The lift command value relative to the operation amount of the operation input unit may also vary depending on the operation range of the operation input unit, the operation speed of the operation input unit, the altitude of the fuselage 16, etc. For example, the altitude of the fuselage 16 may be estimated based on the distance between the ground and the fuselage 16 detected by a ground rangefinder (not shown). For example, the altitude of the fuselage 16 may be estimated based on signals received from a GNSS (Global Navigation Satellite System).

[0056] Even without the pilot inputting any commands to the input unit, the lift command value can be automatically determined for hovering, regardless of the input unit's input level. Furthermore, in the case of automatic control of the aircraft 10, the lift command value can be automatically determined based on a pre-set flight path, also independent of the input unit's input level.

[0057] The roll torque command value calculation unit 56 calculates the roll torque command value. The roll torque command value is determined based on the amount of input by the pilot to the operation input unit. The operation amount from the operation input unit and the roll torque command value may not be a one-to-one correspondence. The roll torque command value corresponding to the operation amount from the operation input unit may also vary depending on the operating range of the operation input unit, the operating speed of the operation input unit, the altitude of the fuselage 16, etc. The angular velocity of the fuselage 16 is detected, for example, by a gyroscope sensor (not shown).

[0058] Even without the pilot inputting any commands to the input unit, the roll torque command can be automatically determined for hovering, regardless of the input unit's input. Furthermore, when the aircraft 10 is automatically controlled, the roll torque command value can be automatically determined based on a pre-set flight path, also independent of the input unit's input.

[0059] The pitch moment command calculation unit 58 calculates the pitch moment command value. For example, it determines the pitch moment command value based on the pilot's input to the control input unit. However, the input to the control input unit and the pitch moment command value may not be a one-to-one correspondence. The pitch moment command value corresponding to the input to the control input unit may also vary depending on the operating range of the control input unit, the operating speed of the control input unit, the altitude of the fuselage 16, etc.

[0060] Even without the pilot inputting any commands to the input unit, the system automatically determines the pitch moment command and hovers, regardless of the input input. Furthermore, when the aircraft is automatically controlled, the system can also automatically determine the pitch moment command based on a pre-set flight path, regardless of the input input.

[0061] The yaw moment command calculation unit 60 calculates the yaw moment command value. For example, the yaw moment command value is determined based on the pilot's input to the control input unit. The input to the control input unit and the yaw moment command value may not be a one-to-one correspondence. The yaw moment command value corresponding to the input to the control input unit may also vary depending on the control input unit's operating range, operating speed, fuselage height, etc.

[0062] Even without the pilot inputting any commands to the input unit, the yaw moment command is automatically determined, and hovering is initiated, regardless of the input unit's input. Furthermore, when the aircraft is automatically controlled, the yaw moment command can also be automatically determined based on a pre-set flight path, independent of the input unit's input.

[0063] The mode selection unit 62 selects between normal mode and energy-saving mode. The selection of normal mode or energy-saving mode can also be based on the pilot's operation of a switch (not shown). Alternatively, the selection can be automatic, based on factors such as the remaining fuel level.

[0064] In the normal mode, such as during cruise, the lift and torque acting on the fuselage 16 are controlled by adjusting the rotation stop angle of the VTOL rotor 12 and the pitch of the blades 26. In the energy-saving mode, such as during cruise, the drag acting on the fuselage 16 is reduced compared to the normal mode by adjusting the rotation stop angle of the VTOL rotor 12 and the pitch of the blades 26. Furthermore, even when the lift and torque acting on the fuselage 16 are controlled by adjusting the rotation stop angle of the VTOL rotor 12 and the pitch of the blades 26, the lift and torque acting on the fuselage 16 can also be controlled by adjusting the control surfaces of the canard wing 18 and the aft wing 20.

[0065] The thrust command calculation unit 64 calculates the thrust command value. For example, the thrust command value is determined based on the amount of input by the pilot to the operation input unit. The amount of input and the thrust command value may not be in a one-to-one correspondence. The thrust command value corresponding to the amount of input can also vary depending on the operating range of the operation input unit, the operating speed of the operation input unit, the altitude of the fuselage 16, etc. For example, the airspeed of the fuselage 16 is detected by an airspeed sensor (not shown).

[0066] The thrust command value can be automatically determined even without the pilot inputting any commands to the control input unit, allowing the aircraft 10 to fly at a constant speed. Furthermore, when the aircraft 10 is automatically controlled, the thrust command value can be automatically determined based on a pre-set flight path, regardless of the input input unit.

[0067] The allocation command value calculation unit 66 calculates allocation command values ​​for lift command values, roll moment command values, pitch moment command values, and yaw moment command values, respectively. The allocation command value represents the command value allocated to the control surfaces of the canard 18 and aft 20, and the command value allocated to the control of the VTOL rotor 12. The allocation command value can be calculated based on a mode selected from normal mode and energy-saving mode. Alternatively, the allocation command value can be calculated based on the magnitude of each command value.

[0068] The control surface control command value generation unit 68 generates drive command values ​​for the aileron, elevator, and rudder based on the control surface control allocation command values ​​assigned to the canard wing 18 and aft wing 20. The control surface control unit 70 controls the drive mechanisms that drive the aileron, elevator, and rudder based on these drive command values.

[0069] When the mode selection unit 62 selects the normal mode, the VTOL rotor command value generation unit 72 generates the drive command value for the VTOL motor 30 based on the control allocation command value assigned to the VTOL rotor 12. The VTOL rotor command value generation unit 72 also generates the drive command value for the blades 26 of the VTOL rotor 12 based on the control allocation command value assigned to the VTOL rotor 12.

[0070] When the mode selector 62 selects the energy-saving mode, the VTOL rotor command value generation unit 72 sets the rotation stop position of the VTOL motor 30 and the pitch of the blade 26 to reduce the drag generated at the VTOL rotor 12 compared to the normal mode. Based on the set rotation stop position of the VTOL motor 30 and the pitch of the blade 26, the VTOL rotor command value generation unit 72 generates drive command values ​​for the VTOL motor 30 and the pitch drive mechanism 34. The VTOL rotor command value generation unit 72 can set the rotation stop position of the VTOL motor 30 and the pitch of the blade 26 to minimize the drag generated at the VTOL rotor 12.

[0071] The VTOL motor control unit 74 controls the inverter 32 according to the drive command of the VTOL motor 30. The pitch control unit 76 controls the pitch drive mechanism 34 according to the drive command value of the blade 26 of the VTOL rotor 12.

[0072] The cruise rotor command value generation unit 78 generates drive command values ​​for the cruise motor 40 based on the thrust command values. The cruise rotor command value generation unit 78 also generates drive command values ​​for the blades of the cruise rotor 14 based on the thrust command values. The cruise motor control unit 80 controls the inverter 42 based on the drive commands from the cruise motor 40. The pitch control unit 82 controls the pitch drive mechanism 44 based on the drive command values ​​from the blades of the cruise rotor 14.

[0073] In the above description, the thrust command value calculated in the thrust command value calculation unit 64 is directly input to the cruise rotor command value generation unit 78. The cruise rotor command value generation unit 78 generates the drive command value for the cruise motor 40 based on the thrust command value.

[0074] In contrast, the thrust command value can also be input to the allocation command value calculation unit 66 in the same way as the lift command value, roll moment command value, pitch moment command value, and yaw moment command value. In this case, the allocation command value calculation unit 66 calculates the allocation command value for each of the lift command value, roll moment command value, pitch moment command value, yaw moment command value, and thrust command value. The cruise rotor command value generation unit 78 can also generate the drive command value for the cruise motor 40 based on the command value allocated to the cruise rotor 14.

[0075] [Regarding the forces generated in the VTOL rotor when rotation stops] As described above, in the aircraft 10 of this embodiment, when the VTOL rotor 12 is in a stopped state, the stop angle of the VTOL rotor 12 and the pitch of the blade 26 are controlled. Accordingly, the magnitude of the lift generated by each VTOL rotor 12 is controlled. As a result, the magnitude of the lift acting on the fuselage 16 and the magnitude of the torque acting around the center of gravity G of the fuselage 16 are controlled.

[0076] When the torque acts around the center of gravity G of the fuselage 16, the control device 48 ensures that the torque generated by the force at the outer rotor, located away from the center of gravity G, is greater than the torque generated by the force at the inner rotor, located closer to the center of gravity G. When the forces generated at the outer rotor and the inner rotor are the same, the torque generated by the force at the outer rotor acting around the center of gravity G of the fuselage 16 is greater than the torque generated by the force at the inner rotor acting around the center of gravity G of the fuselage 16. Therefore, the control device 48 can suppress the forces generated at both the outer and inner rotors and ensure the torque acting around the center of gravity G of the fuselage 16.

[0077] When the rolling moment acts around the center of gravity G of the fuselage 16, the control device 48 creates a difference between the total lift of the four VTOL rotors 12 located to the left of the center of gravity G and the total lift of the four VTOL rotors 12 located to the right of the center of gravity G. In this case, the control device 48 ensures that the magnitude of the rolling moment acting around the center of gravity G of the fuselage 16 due to the lift generated by the outer rotors is greater than the magnitude of the rolling moment acting around the center of gravity G of the fuselage 16 due to the lift generated by the inner rotors. At this time, the control device 48 can make the total lift generated by each outer rotor greater than the total lift generated by each inner rotor.

[0078] When the pitching moment acts around the center of gravity G of the fuselage 16, the control device 48 creates a difference between the total lift of the four VTOL rotors 12 positioned in front of the center of gravity G and the total lift of the four VTOL rotors 12 positioned behind the center of gravity G. In this case, the control device 48 ensures that the pitching moment acting around the center of gravity G of the fuselage 16 due to the lift generated by the outer rotors is greater than the pitching moment acting around the center of gravity G of the fuselage 16 due to the lift generated by the inner rotors. At this time, the control device 48 can make the total lift generated by each outer rotor greater than the total lift generated by each inner rotor.

[0079] When the yaw moment acts around the center of gravity G of the fuselage 16, the control device 48 causes a difference between the total drag of the four VTOL rotors 12 located to the left of the center of gravity G and the total drag of the four VTOL rotors 12 located to the right of the center of gravity G. In this case, the control device 48 causes the magnitude of the yaw moment acting around the center of gravity G of the fuselage 16 due to the drag generated by the outer rotors to be greater than the magnitude of the yaw moment acting around the center of gravity G of the fuselage 16 due to the drag generated by the inner rotors. At this time, the control device 48 can make the total lift generated at each outer rotor greater than the total lift generated at each inner rotor.

[0080] [VTOL rotor rotation stop angle control and blade pitch control] When the VTOL rotor 12 is in a stopped state, the lift and drag generated by each VTOL rotor 12 are controlled by the stopped position of the VTOL rotor 12 and the pitch of the blade 26.

[0081] When the lift and drag generated by each VTOL rotor 12 change to the command value, the control device 48 controls the pitch of the blade 26 to change the lift and drag, and then controls the rotation stop position of the VTOL rotor 12 to change the lift and drag.

[0082] Inverter 32 controls VTOL motor 30, outputting torque to VTOL motor 30 that balances the external force, thereby maintaining the rotational stop position of VTOL rotor 12. When changing the rotational stop position of VTOL rotor 12, it is necessary to break the balance between the external force and the torque of VTOL motor 30. Compared to breaking the balance between the external force and the torque of VTOL motor 30, changing the pitch of blade 26 has a faster response speed. Therefore, it is possible to improve the response speed to changes in lift and drag generated by each VTOL rotor 12.

[0083] [Effects] In the aircraft 10 of this embodiment, when the fuselage 16 has an airspeed, lift and drag can be generated in the VTOL rotor 12 when the VTOL rotor 12 is in a state of rotational stop.

[0084] Furthermore, in the aircraft 10 of this embodiment, the control device 48 controls the stop rotation angle of the VTOL rotor 12 and the pitch of the blades 26, and controls the force generated by each VTOL rotor 12 when the VTOL rotor 12 is in a stopped state, thereby causing the force to act on the fuselage 16. Therefore, even when the VTOL rotor 12 is in a stopped state, the force generated by each VTOL rotor 12 can be used to cause the force to act on the fuselage 16.

[0085] Furthermore, in the aircraft 10 of this embodiment, when a torque is generated around the center of gravity G of the fuselage 16, the control device 48 ensures that the torque exerted by the force generated by the outer rotor around the center of gravity G of the fuselage 16 is greater than the torque exerted by the force generated by the inner rotor around the center of gravity G of the fuselage 16. Therefore, the control device 48 can suppress the forces generated at the outer and inner rotors and ensure the torque exerted around the center of gravity G of the fuselage 16.

[0086] Furthermore, in the aircraft 10 of this embodiment, the control device 48 controls the pitch of the rotor blades 26 to change lift and drag, and then controls the stop position of the VTOL rotor 12 to change lift and drag. Accordingly, the control device 48 can improve the response speed when the lift and drag generated at each VTOL rotor 12 change.

[0087] Furthermore, in the aircraft 10 of this embodiment, when an energy-saving mode that prioritizes reducing drag on the fuselage 16 is selected, the control device 48 controls the rotation stop angle of each VTOL rotor 12 and the pitch of each blade 26, thereby reducing the drag generated on each VTOL rotor 12 compared to other modes. Accordingly, the aircraft 10 can reduce energy consumption.

[0088] [Variation Example] The VTOL rotor 12 of the aircraft 10 in the first embodiment generates lift that moves the fuselage 16 upward by rotating. Alternatively, a tilting rotor capable of changing the orientation of the rotor relative to the fuselage 16 may be used instead of the VTOL rotor 12. When the tilting rotor is oriented at a predetermined angle relative to the fuselage 16, the tilting rotor generates lift that moves the fuselage 16 upward by rotating.

[0089] In the first embodiment, the aircraft 10 has multiple (four in the first embodiment) VTOL rotors 12 arranged in the longitudinal direction of the fuselage 16. Alternatively, multiple VTOL rotors 12 of the aircraft 10 can be arranged in the lateral direction of the fuselage 16. Furthermore, multiple VTOL rotors 12 of the aircraft 10 can also be arranged in the vertical direction of the fuselage 16.

[0090] In the first embodiment, the VTOL rotor 12 is connected to the VTOL motor 30. A gearbox can be disposed between the VTOL motor 30 and the VTOL rotor 12. The gearbox is a transmission used to change the rotational speed of the VTOL motor 30 and transmit it to the VTOL rotor 12. Similarly, a gearbox can also be disposed between the cruise motor 40 and the cruise rotor 14.

[0091] [Invention derived from implementation methods] The invention described below can be understood from the above embodiments.

[0092] A control device (48) for an aircraft (10) having a plurality of rotors (12) and one or more fixed wings (18, 20), wherein each of the plurality of rotors has one or more blades (26) and generates lift by rotating about a rotation axis (24); the one or more fixed wings (18, 20) generate lift when the fuselage (16) has an airspeed, and the control device controls the force generated by each rotor when the rotors are stopped by controlling the pitch of each blade and the rotation stop angle of each rotor, thereby causing the force to act on the fuselage. Therefore, even when the rotors are stopped, the force generated by each rotor can be used to act on the fuselage.

[0093] In the aforementioned aircraft control device, the aircraft may have one or more inner rotors (12FLb, 12RLb, 12FRb, 12RRb) and one or more outer rotors (12FLa, 12RLa, 12FRa, 12RRa) as the rotors. The distance from the center of gravity (G) of the fuselage to the inner rotor is shorter than the distance from the center of gravity to the outer rotor. When the control device generates a torque around the center of gravity of the fuselage, it ensures that the torque acting on the fuselage through the force generated by the outer rotor is greater than the torque acting on the fuselage through the force generated by the inner rotor. Accordingly, the control device can reduce the force generated on the rotor relative to the required torque magnitude.

[0094] In the aforementioned aircraft control device, the aircraft may also have a canard (18), a rear wing (20), one or more first rotors (12FLa, 12FRa), one or more second rotors (12FLb, 12FRb), one or more third rotors (12RLb, 12RRb), and one or more fourth rotors (12RLa, 12RRa), wherein the canard serves as the fixed wing and is positioned forward of the fuselage's center of gravity; the rear wing serves as the fixed wing and is positioned rearward of the fuselage's center of gravity; the one or more first rotors serve as the outer rotors and are positioned forward of the canard; the one or more second rotors serve as the inner rotors and are positioned between the canard and the center of gravity in the longitudinal direction of the fuselage; the one or more third rotors serve as the inner rotors and are positioned between the center of gravity and the rear wing in the longitudinal direction of the fuselage; and the one or more fourth rotors serve as the outer rotors and are positioned rearward of the rear wing. Accordingly, the control device can reduce the force generated on the rotor relative to the required torque.

[0095] In the aforementioned aircraft control device, when the force generated by each rotor changes, the control device can change the force generated by each rotor by controlling the pitch of each blade, and then change the force generated by each rotor by controlling the rotation stop angle of each rotor. Accordingly, the control device can improve its response speed when the force generated by each rotor changes.

[0096] In the aforementioned aircraft control system, when an energy-saving mode that prioritizes reducing drag on the fuselage is selected, the rotation stop angle of each rotor and the pitch of each blade can be controlled to reduce drag generated on each rotor compared to selecting other modes. Accordingly, the aircraft can reduce energy consumption.

[0097] Furthermore, the present invention is not limited to the above disclosure, and various structures can be adopted without departing from the spirit of the present invention.

Claims

1. A control device (48) for an aircraft (10) having a plurality of rotors (12) and one or more fixed wings (18, 20), wherein, Each of the plurality of rotors has one or more blades (26) and generates lift by rotating about a rotation axis (24); the one or more fixed wings (18, 20) generate lift when the fuselage (16) has an airspeed. Its features are, When the fixed wing generates lift, the lift and drag generated by each rotor when the rotor rotation is stopped are controlled by controlling the pitch of each rotor blade and the stop angle of each rotor. This generates torque around the center of gravity of the fuselage, thereby controlling the attitude of the aircraft.

2. The aircraft control device according to claim 1, characterized in that, The aircraft has one or more internal rotors (12FLb, 12RLb, 12FRb, 12RRb) as the rotor and one or more external rotors (12FLa, 12RLa, 12FRa, 12RRa) as the rotor. The distance from the center of gravity (G) of the fuselage to the inner rotor is shorter than the distance from the center of gravity to the outer rotor. When a torque is generated around the center of gravity of the fuselage, the torque acting on the fuselage through the force generated by the outer rotor is greater than the torque acting on the fuselage through the force generated by the inner rotor.

3. The aircraft control device according to claim 2, characterized in that, The aircraft has a canard (18), a rear wing (20), one or more first rotors (12FLa, 12FRa), one or more second rotors (12FLb, 12FRb), one or more third rotors (12RLb, 12RRb), and one or more fourth rotors (12RLa, 12RRa), wherein, The forewing, which serves as the fixed wing, is positioned forward of the fuselage's center of gravity. The rear wing, as the fixed wing, is positioned further rearward than the center of gravity of the fuselage. The one or more first rotors, as the outer rotors, are positioned forward of the forewing; The one or more second rotors serve as the inner rotors and are positioned between the forewing and the center of gravity in the longitudinal direction of the fuselage; The one or more third rotors, as the inner rotors, are positioned in the longitudinal direction of the fuselage between the center of gravity and the rear wing; The one or more fourth rotors, as the outer rotors, are positioned further rearward than the rear wing.

4. The control device for an aircraft according to any one of claims 1 to 3, characterized in that, When the force generated by each rotor changes, the force generated by each rotor is changed by controlling the pitch of each blade, and then the force generated by each rotor is changed by controlling the rotation stop angle of each rotor.

5. The control device for an aircraft according to any one of claims 1 to 3, characterized in that, When the energy-saving mode that prioritizes reducing drag on the fuselage is selected, the rotation stop angle of each rotor and the pitch of each blade are controlled to reduce the drag generated on each rotor compared to selecting other modes.

Images