Hybrid aircraft

Abstract

The present invention discloses a hybrid aircraft. The hybrid aircraft (10) uses a plurality of rotors (a takeoff and landing rotor (20), a cruise rotor (22)) for flight. The hybrid aircraft (10) includes a generator (MG (76)), a battery (78), a plurality of electric motors (68), and a controller (64). Among them, the battery (78) is charged with the electric power generated by the generator; the plurality of electric motors (68) rotate each rotor by the electric power generated by the generator and the electric power supplied from the battery (78); the controller (64) sets a target remaining capacity (target (SOC)) of the battery (78) according to the flight state of the aircraft body, and controls the charge and discharge of the battery (78) so that the remaining capacity (SOC) of the battery (78) approaches the target remaining capacity. Accordingly, overcharging and undercharging of the battery can be prevented.

Description

Technical Field

[0001] The present invention relates to a hybrid aircraft that supplies electric power from a generator and a battery to an electric motor to rotate a rotor. Background Art

[0002] An aircraft called an electric vertical take-off and landing aircraft (eVTOL aircraft) is shown in the specification of US Patent Application Publication No. 2016 / 0023773. This aircraft is a so-called hybrid aircraft. The hybrid aircraft is configured to be able to supply electric power from a generator driven by a gas turbine engine and an energy storage system (such as a battery) to an electric motor to rotate a rotor. The hybrid aircraft flies mainly by supplying electric power generated by the generator to the electric motor. One flight of the aircraft includes take-off, ascent, cruise, approach, and landing. In the specification of US Patent Application Publication No. 2016 / 0023773, the maximum electric power required for take-off, ascent, approach, and landing is shown. Hereinafter, take-off, ascent, approach, and landing are sometimes collectively referred to as take-off and landing. Summary of the Invention

[0003] When a disturbance such as a strong gust occurs during the flight of a hybrid aircraft, the attitude of the aircraft sometimes loses balance. In this case, the controller of the hybrid aircraft controls each rotor to re-align the attitude of the aircraft. At this time, the controller repeatedly changes the requested value of the electric power within a short period of time. The requested value of the electric power is referred to as the electric power request value. It is difficult to make the generated power follow the change of the electric power request value. Therefore, when the generated power exceeds the electric power request value, the excess electric power is absorbed by charging the battery. In addition, when the generated power is less than the electric power request value, the insufficient electric power is supplemented by discharging the battery.

[0004] The electric power request value during take-off and landing is large. When a disturbance occurs during take-off and landing, the variation range of the electric power request value becomes large. At this time, the excess or insufficient electric power also becomes large, and the charge and discharge amount of the battery becomes large. Generally, the controller controls the generated power so that the remaining capacity of the battery (SOC: State Of Charge) approaches the target value (target SOC). When the target SOC is set relatively high, overcharging may occur when the battery is charged with excess electric power. Overcharging causes deterioration of the battery, so it is best not to occur overcharging. On the other hand, when the target SOC is set relatively low, it is no problem to charge the battery with excess electric power. However, if the target SOC is set relatively low, the remaining capacity of the battery may become insufficient due to continuous use of the battery.

[0005] An object of the present invention is to solve the above technical problems.

[0006] The method of the present invention is a hybrid aircraft that uses multiple rotors for flight.

[0007] It has a generator, a battery, multiple electric motors, and a controller. Among them,

[0008] the battery is charged by the electric power generated by the generator;

[0009] multiple of the electric motors rotate each of the rotors by the electric power generated by the generator and the electric power supplied by the battery;

[0010] the controller sets a target remaining capacity of the battery according to the flight state of the aircraft body, and controls the charge and discharge of the battery so that the remaining capacity of the battery approaches the target remaining capacity.

[0011] According to the present invention, overcharging and undercharging of the battery can be prevented.

[0012] Based on the description of the following embodiments with reference to the drawings, the above objects, features, and advantages should be easily understood. Description of the Drawings

[0013] Figure 1 is a schematic top view of the hybrid aircraft.

[0014] Figure 2 is a block diagram showing the structure of the electric power system of the hybrid aircraft.

[0015] Figure 3A is a graph showing the change in elapsed time from takeoff to landing and the generated power. Figure 3B is a graph showing the change in elapsed time from takeoff to landing and the electric power request value.

[0016] Figure 4 is a graph showing Figure 3A the generated power of Figure 3B and the change in the difference amount of the electric power request value of

[0017] Figure 5 is a block diagram showing the processing procedure of the hybrid ECU.

[0018] Figure 6 is a graph showing the SOC conversion table.

[0019] FIG. 7A is a graph showing the passage of time and the change in altitude from before takeoff to after landing of the hybrid aircraft. FIG. 7B is a graph showing the passage of time and the stable output of the battery from before takeoff to after landing of the hybrid aircraft. FIG. 7C is a graph showing the passage of time and the change in the electric power variation value from before takeoff to after landing of the hybrid aircraft. FIG. 7D is a graph showing the passage of time and the change in SOC and the change in the target SOC from before takeoff to after landing of the hybrid aircraft.

[0020] FIG. 8A is a graph showing the passage of time and the change in altitude from before takeoff to after landing of the hybrid aircraft. FIG. 8B is a graph showing the passage of time and the stable output of the battery from before takeoff to after landing of the hybrid aircraft. FIG. 8C is a graph showing the passage of time and the change in the electric power variation value from before takeoff to after landing of the hybrid aircraft. FIG. 8D is a graph showing the passage of time and the change in SOC and the change in the target SOC from before takeoff to after landing of the hybrid aircraft. DETAILED DESCRIPTION

[0021] [1 Overall Structure of Hybrid Aircraft 10]

[0022] Use Figure 1 to describe the overall structure of the hybrid aircraft 10. In this specification, the hybrid aircraft 10 is also simply referred to as the aircraft 10. The aircraft 10 is an electric vertical takeoff and landing aircraft. The electric vertical takeoff and landing aircraft is called an eVTOL aircraft. The eVTOL aircraft generates lift and thrust respectively by rotors with an electric motor 68 ( Figure 2 ). Electric power generated by power generation and stored electric power are respectively supplied to the electric motor 68. In addition, in this specification, the vertically upward direction is defined as the upper side. The vertically downward direction is defined as the lower side. In addition, when the aircraft 10 moves (flies) in the horizontal direction, the moving direction of the aircraft 10 is defined as the front. The opposite direction of the front is defined as the rear. In addition, when observing the aircraft 10 moving forward, the right side is defined as the right side and the left side is defined as the left side. In addition, looking down at the aircraft 10 refers to the state of observing the aircraft 10 from above. Observing the aircraft 10 from the front refers to the state of observing the aircraft 10 from the front.

[0023] The aircraft 10 includes a fuselage 12, a front wing 14, a rear wing 16, two booms 18, eight takeoff and landing rotors 20, and two cruise rotors 22. The central axis A of the fuselage 12 extends in the front-rear direction. The structure of the aircraft 10 is symmetric about a vertical plane including the central axis A. When observed from above, the central axis A coincides with the center of gravity G of the aircraft 10.

[0024] The fuselage 12 is long in the front-rear direction. The fuselage 12 has a front fuselage part 12f and a rear fuselage part 12r. The front fuselage part 12f is located in front of the center of gravity G. The rear fuselage part 12r is located behind the center of gravity G. The front fuselage part 12f tapers as it approaches the front end. The rear fuselage part 12r tapers as it approaches the rear end. Additionally, the fuselage 12 has a main body. The fuselage 12 may also have a main body and a fairing that covers a part of the main body. In this specification, the main body and the fairing are referred to as the fuselage 12. The front part in the fuselage 12 is called the front fuselage part 12f. The rear part in the fuselage 12 is called the rear fuselage part 12r.

[0025] The front wing 14 is connected to the upper part of the front fuselage part 12f. The front wing 14 generates lift when the aircraft 10 moves forward.

[0026] The rear wing 16 is connected to the upper part of the rear fuselage part 12r. The rear wing 16 generates lift when the aircraft 10 moves forward.

[0027] The wing area of the rear wing 16 is larger than the wing area of the front wing 14. Additionally, the wing width of the rear wing 16 is longer than the wing width of the front wing 14. With this structure, the lift generated by the rear wing 16 when the aircraft 10 moves forward is greater than the lift generated by the front wing 14. That is, the rear wing 16 functions as the main wing of the aircraft 10. The rear wing 16 is a swept wing that reduces air resistance. On the other hand, the front wing 14 functions as a forewing of the aircraft 10.

[0028] Alternatively, it can be such that the lift generated by the rear wing 16 when the aircraft 10 moves forward and the lift generated by the front wing 14 when the aircraft 10 moves forward are of the same degree. The ratio of the lift generated by the front wing 14 and the lift generated by the rear wing 16 is appropriately determined according to the position of the center of gravity G, the posture of the aircraft during cruising, etc. Additionally, the sizes of the front wing 14 and the rear wing 16 are determined to generate the desired lift. In this specification, the size of the wing is the wing area, length, etc.

[0029] The two cantilevers 18 include a right cantilever 18R and a left cantilever 18L. The right cantilever 18 is arranged on the right side of the fuselage 12. The left cantilever 18 is arranged on the left side of the fuselage 12. The two cantilevers 18 form a pair. The two cantilevers 18 are arranged symmetrically left and right with respect to the vertical plane including the central axis A. The two cantilevers 18 are connected to the front wing 14 and the rear wing 16. The two cantilevers 18 are connected to the fuselage 12 through the front wing 14 and the rear wing 16. The two cantilevers 18 each function as a support member for supporting four landing and takeoff rotors 20.

[0030] The cantilever 18R on the right side is a rod-shaped member. The cantilever 18R on the right side extends from the front to the rear. The cantilever 18R on the right side is connected to the wing tip on the right side of the front wing 14. The cantilever 18R on the right side is connected to the rear wing 16. The front end of the cantilever 18R on the right side is located in front of the front wing 14. The rear end of the cantilever 18R on the right side is located behind the rear wing 16.

[0031] The cantilever 18L on the left side is a rod-shaped member. The cantilever 18L on the left side extends from the front to the rear. The cantilever 18L on the left side is connected to the wing tip on the left side of the front wing 14. The cantilever 18L on the left side is connected to the rear wing 16. The front end of the cantilever 18L on the left side is located in front of the front wing 14. The rear end of the cantilever 18L on the left side is located behind the rear wing 16.

[0032] Each of the eight takeoff and landing rotors 20 has a mast (not shown), a hub (not shown), and a plurality of blades 46. The mast is connected to the output shaft portion of the electric motor 68. The hub is connected to the mast. The plurality of blades 46 are connected to the hub. The mast is arranged parallel to the vertical direction. The mast can rotate about a rotation axis extending in the vertical direction. The plurality of blades 46 are arranged above the cantilevers 18, the front wing 14, and the rear wing 16. The pitch angle of the blades 46 is adjustable. With this structure, the takeoff and landing rotor 20 rotates about the rotation axis to generate lift. One rotor unit for generating lift has one takeoff and landing rotor 20, a rotation mechanism (such as the electric motor 68), and a drive circuit. In addition, one rotor unit may have one or more batteries 78.

[0033] The eight takeoff and landing rotors 20 include four takeoff and landing rotors 20a to 20d on the right side and four takeoff and landing rotors 20a to 20d on the left side. The takeoff and landing rotors 20a to 20d on the right side are arranged on the right side of the fuselage 12. The takeoff and landing rotors 20a to 20d on the left side are arranged on the left side of the fuselage 12. The takeoff and landing rotors 20a to 20d on the right side are supported by the cantilever 18R on the right side. The takeoff and landing rotors 20a to 20d on the left side are supported by the cantilever 18L on the left side. The takeoff and landing rotor 20a on the right side and the takeoff and landing rotor 20a on the left side form a pair. The positions of the takeoff and landing rotor 20a on the right side and the takeoff and landing rotor 20a on the left side in the front-rear direction are the same positions. The same applies to the left and right takeoff and landing rotors 20b to 20d.

[0034] As Figure 1 shown, a pair of takeoff and landing rotors 20a, the front wing 14, a pair of takeoff and landing rotors 20b, a pair of takeoff and landing rotors 20c, the rear wing 16, and a pair of takeoff and landing rotors 20d are arranged in order from the front to the rear.

[0035] The two cruising rotors 22 each have a mast (not shown), a hub (not shown), and a plurality of blades (not shown). The mast is connected to the output shaft portion of the electric motor 68. The hub is connected to the mast. The plurality of blades are connected to the hub. A cylindrical duct 54 is provided around the cruising rotor 22. The mast is disposed below the rear wing 16. The mast is arranged parallel to the front-rear direction. The mast can rotate about a rotation axis extending in the front-rear direction. With this structure, the cruising rotor 22 rotates about a rotation axis extending in the front-rear direction to generate thrust. One rotor unit for generating thrust includes one cruising rotor 22, a rotation mechanism (such as the electric motor 68), and a drive circuit. In addition, one rotor unit may have one or more batteries 78.

[0036] The two cruising rotors 22 are respectively disposed at the rear part 12r of the fuselage. The two cruising rotors 22 are located on the left side of the right takeoff / landing rotors 20a to 20d and on the right side of the left takeoff / landing rotors 20a to 20d. In addition, the two cruising rotors 22 are located between the pair of takeoff / landing rotors 20c and the pair of takeoff / landing rotors 20d. In addition, the rotation axes of the two cruising rotors 22 are located below the blades 46 of each of the eight takeoff / landing rotors 20.

[0037] The positions of the two cruising rotors 22 in the front-rear direction are the same as each other. The positions of the two cruising rotors 22 in the up-down direction are also the same as each other. In addition, the two cruising rotors 22 are arranged side by side. The right cruising rotor 22 is disposed at a position on the right side of the vertical plane including the central axis A of the fuselage 12. The right cruising rotor 22 is supported by the right wing of the rear wing 16. The left cruising rotor 22 is disposed at a position on the left side of the vertical plane including the central axis A of the fuselage 12. The left cruising rotor 22 is supported by the left wing of the rear wing 16.

[0038] [Two electric power systems 60]

[0039] Figure 2It is a block diagram showing the structure of the electric power system 60 equipped in the hybrid aircraft 10. The electric power system 60 has a set of gauges 62, a controller 64, a driver 66, an electric motor 68, a fuel control device 70, a gas turbine engine 72, a power control unit 74, a motor generator 76, and a battery 78. In this specification, the gas turbine engine 72 is also referred to as GT72. In this specification, the power control unit 74 is also referred to as PCU74. In this specification, the motor generator 76 is also referred to as MG76. One driver 66 and one electric motor 68 are taken as a group. A group of the driver 66 and the electric motor 68 is provided for each of a plurality of rotors (a plurality of takeoff / landing rotors 20 and a plurality of cruising rotors 22). One fuel control device 70, one GT72, one PCU74, and one MG76 are taken as a group. One or more groups of the group such as GT72 are provided.

[0040] The set of gauges 62 has various gauges for detecting the flight state of the aircraft 10. The set of gauges 62 includes a sensor that measures the operation amount of an operating device operated by a pilot. For example, the sensor measures the operation amount of a joystick or the like. In addition, the set of gauges 62 includes a gauge that measures the rotational speed of the electric motor 68. In addition, the set of gauges 62 may include a gauge that measures the moving speed of the aircraft 10 in the horizontal direction. In addition, the set of gauges 62 may include a gauge that measures the altitude of the aircraft 10. In addition, the set of gauges 62 may include a gauge that measures at least one of the deflection, pitch, and roll of the airframe. In addition, the set of gauges 62 may include a gauge that measures the wind speed and wind direction. Each gauge of the set of gauges 62 periodically sends the measured value to the flight controller 80 of the controller 64.

[0041] The controller 64 has one or more control units. In this embodiment, the controller 64 has a flight controller 80, a hybrid ECU 82, a motor ECU 84, a GT-ECU 86, a PCU-ECU 88, and a BAT-ECU 90. The controller 64 controls the operations of the respective rotors (the takeoff / landing rotors 20 and the cruising rotors 22). The controller 64 controls the SOC of the battery 78 according to the flight state of the aircraft 10.

[0042] The flight controller 80 includes an arithmetic unit 92, a storage unit 94, an input / output unit (not shown), etc. The arithmetic unit 92 includes a processor such as a CPU. The arithmetic unit 92 performs various processes by executing programs stored in the storage unit 94. Additionally, the arithmetic unit 92 may include integrated circuits such as ASICs and FPGAs. Further, the arithmetic unit 92 may include an electronic circuit including discrete devices. The storage unit 94 includes a volatile memory and a non-volatile memory. Examples of the volatile memory include RAM, etc. Examples of the non-volatile memory include ROM, flash memory, etc. The non-volatile memory stores programs executed by the arithmetic unit 92. The non-volatile memory stores various information pre-input by the user.

[0043] The flight controller 80 acquires each measured value from the measuring instrument group 62 and performs various operations and various determinations using a specified algorithm or table. For example, the flight controller 80 determines the flight mode at that time point. The flight mode includes various modes from takeoff to landing. For example, the flight mode includes a vertical takeoff mode M1, a rising mode M2, a cruising mode (horizontal flight mode) M3, a descending mode M4, and a vertical landing mode M5. Additionally, the flight controller 80 calculates the rotational speed of each electric motor 68 required at that time point according to the flight mode and the operation of the operating device. After the calculation, the flight controller 80 outputs a rotational speed command for each electric motor 68 to the motor ECU 84. Additionally, the flight controller 80 calculates the electric power required by the aircraft 10. After the calculation, the flight controller 80 outputs an electric power request value [unit: w] to the hybrid ECU 82. Additionally, the flight controller 80 determines the electric power change value [unit: w] corresponding to the flight mode. After the determination, the flight controller 80 outputs the electric power change value to the hybrid ECU 82.

[0044] Here, the flight mode, the electric power request value, and the electric power change value will be described. Figure 3A It is a graph showing the change in the elapsed time from takeoff to landing and the power generation. Figure 3B It is a graph showing the change in the elapsed time from takeoff to landing and the electric power request value. Figure 4 Indicates Figure 3A The power generation of Figure 3B And the difference amount (electric power difference amount) of the electric power request value of Figure 4 In Figure 4 The shown electric power difference amount periodically repeats going up and down with respect to the power generation (zero on the vertical axis). The maximum change amount of the electric power difference amount per unit time is called the electric power change value. For example, the electric power change value corresponds to the change amount from the maximum value on the positive side to the maximum value on the negative side in the electric power difference amount of one cycle. The electric power change value is approximately a constant value for each flight mode.

[0045] As Figure 3B shown, the electric power request value is the largest when the flight mode is the vertical takeoff mode M1. The electric power request value gradually decreases as the flight mode transfers to the ascent mode M2 and the cruise mode M3. The electric power request value reaches the minimum in the descent mode M4. Also, the electric power request value reaches the maximum again when the flight mode is the vertical landing mode M5. The electric power variation value is roughly proportional to the magnitude of the electric power request value. Therefore, as Figure 4 shown, the electric power variation value also reaches the maximum in the vertical takeoff mode M1 and the vertical landing mode M5.

[0046] The takeoff and landing rotor 20 is used in the vertical takeoff mode M1 and the vertical landing mode M5. In the ascent mode M2, the takeoff and landing rotor 20 and the cruise rotor 22 are used in combination. In the ascent mode M2, the usage rate of the takeoff and landing rotor 20 gradually decreases, and the usage rate of the cruise rotor 22 gradually increases. The cruise rotor 22 is used in the cruise mode M3. In the descent mode M4, the takeoff and landing rotor 20 and the cruise rotor 22 are used in combination. In the descent mode M4, the usage rate of the takeoff and landing rotor 20 gradually increases, and the usage rate of the cruise rotor 22 gradually decreases. The electric power request value and the electric power variation value become large in the vertical takeoff mode M1 and the vertical landing mode M5. The storage unit 94 stores in advance the relationship among the flight mode, the electric power request value, and the electric power variation value. The arithmetic unit 92 determines the flight mode and calculates the electric power request value and the electric power variation value.

[0047] Return Figure 2 Continue to describe the electric power system 60. The hybrid ECU 82 includes an arithmetic unit 96, a storage unit 98, an input / output unit (not shown), etc. The arithmetic unit 96 is the same as the arithmetic unit 92 and is composed of a processor such as a CPU, for example. The storage unit 98 is the same as the storage unit 94 and is composed of a volatile memory and a non-volatile memory, for example. The non-volatile memory stores the SOC conversion table TB( Figure 6 ).

[0048] The hybrid ECU 82 acquires the electric power request value and the electric power variation value from the flight controller 80. The hybrid ECU 82 acquires the SOC of the battery 78 from the BAT-ECU 90. Based on the acquired information, the hybrid ECU 82 calculates the required power generation amount at that time point. After the calculation, the hybrid ECU 82 outputs the rotational speed command and the torque command for each GT 72 to the GT-ECU 86. After the calculation, the hybrid ECU 82 outputs the rotational speed command and the torque command for each MG 76 to the PCU-ECU 88. The arithmetic process of the hybrid ECU 82 will be described in [3] below.

[0049] The structure of the motor ECU 84 is also the same as that of the hybrid ECU 82. The motor ECU 84 obtains the rotational speed command from the flight controller 80 and outputs control signals to the respective drivers 66.

[0050] The structure of the GT-ECU 86 is also the same as that of the hybrid ECU 82. The GT-ECU 86 obtains the rotational speed command and torque command from the hybrid ECU 82 and outputs control signals to the respective fuel control devices 70.

[0051] The structure of the PCU-ECU 88 is also the same as that of the hybrid ECU 82. The PCU-ECU 88 obtains the rotational speed command and torque command from the hybrid ECU 82 and outputs control signals to the respective PCUs 74.

[0052] The structure of the BAT-ECU 90 is also the same as that of the hybrid ECU 82. The BAT-ECU 90 obtains information such as the voltage, current, and temperature of the battery 78 from a plurality of sensors (not shown) and calculates the SOC using a prescribed algorithm or table. The BAT-ECU 90 outputs the SOC to the hybrid ECU 82.

[0053] The driver 66 is a drive circuit for the electric motor 68. For example, the driver 66 is a circuit including an inverter or the like. The driver 66 is interposed between the MG 76 and the electric motor 68. The driver 66 is also interposed between the battery 78 and the electric motor 68. The driver 66 supplies electric power to the electric motor 68 from at least one of the MG 76 and the battery 78 according to the control signal output from the motor ECU 84.

[0054] The fuel control device 70 is a device that controls the fuel supplied to the GT 72. For example, the fuel control device 70 is a device including a fuel control valve. The fuel control device 70 supplies fuel to the GT 72 according to the control signal output from the GT-ECU 86.

[0055] The PCU 74 has an output control circuit for the MG 76 and a drive circuit for the MG 76. For example, the PCU 74 has a circuit including a DC / DC converter, an inverter, etc. The PCU 74 is interposed between the MG 76 and the electric motor 68, and between the MG 76 and the battery 78. The PCU 74 operates according to the control signal output from the PCU-ECU 88 to supply electric power from the MG 76 to the electric motor 68 and the battery 78, respectively. In addition, the PCU 74 operates according to the control signal output from the PCU-ECU 88 to supply electric power from the battery 78 to the MG 76.

[0056] A plurality of electric motors 68 are connected to the rotor (the takeoff / landing rotor 20 or the cruise rotor 22). The GT 72 and the MG 76 are connected to each other.

[0057] [3 Hybrid ECU 82]

[0058] Figure 5 is a block diagram showing the processing steps of the hybrid ECU 82. The arithmetic unit 96 of the hybrid ECU 82 functions as an electric energy estimator 102, an SOC calculator 104, a charge / discharge amount calculator 106, an adder 108, and a power generation controller 110.

[0059] The electric energy estimator 102 estimates the electric energy change value [unit: wh] based on the electric power change value obtained from the flight controller 80. The electric energy estimator 102 calculates the electric energy change value by multiplying the electric power change value by Figure 4 the time of one cycle of the shown electric power difference amount.

[0060] The SOC calculator 104 calculates the target SOC based on the electric energy change value obtained from the electric energy estimator 102 and the SOC conversion table TB. As Figure 6 shown, the SOC conversion table TB is set such that the target SOC decreases as the electric energy change value increases.

[0061] The charge / discharge amount calculator 106 calculates the target charge / discharge amount [unit: w] based on the target SOC obtained from the SOC calculator 104 and the SOC obtained from the BAT-ECU 90. The charge / discharge amount calculator 106 performs PID control. That is, the charge / discharge amount calculator 106 calculates the target charge / discharge amount for making the SOC approach the target SOC.

[0062] The adder 108 adds the target charge / discharge amount obtained from the charge / discharge amount calculator 106 and the electric power request value obtained from the flight controller 80 as the target power generation amount [unit: w].

[0063] The power generation controller 110 calculates the rotation speed and torque of the GT72 required for the target power generation amount obtained from the adder 108. The power generation controller 110 also calculates the rotation speed and torque of the MG76 required for the target power generation amount obtained from the adder 108. The power generation controller 110 respectively outputs a rotation speed command for the GT72 and a torque command for the GT72 to the GT-ECU 86. The power generation controller 110 respectively outputs a rotation speed command for the MG76 and a torque command for the MG76 to the PCU-ECU 88.

[0064] [4 Control Example of SOC]

[0065] [4.1 First Control Example]

[0066] The first control example of the SOC from before the aircraft 10 takes off to after it lands is described with reference to FIGS. 7A to 7D. In FIG. 7D, the solid line represents the target SOC, and the dashed line represents the actual SOC. As shown in FIG. 7A, at time point t1, the operator performs the start operation of GT72. At time point t2, the start of GT72 is completed. At time point t3, the aircraft takes off and ascends vertically. At time point t4, the aircraft starts to move forward while ascending. At time point t5, the aircraft starts horizontal flight, i.e., starts cruising. At time point t6, the aircraft starts to descend while moving forward. At time point t7, the aircraft descends vertically. At time point t8, the aircraft lands. At time point t9, the operator stops GT72.

[0067] The flight mode from time point t3 to time point t4 is the vertical takeoff mode M1. The flight mode between time point t4 and time point t5 is the ascending mode M2. The flight mode between time point t5 and time point t6 is the cruising mode M3. The flight mode between time point t6 and time point t7 is the descending mode M4. The flight mode between time point t7 and time point t8 is the vertical landing mode M5.

[0068] Between time point t1 and time point t3, the controller 64 controls the PCU 74 and the driver 66 to supply the electric power of the battery 78 to the MG 76 and the electric motor 68. During this period, as shown in FIG. 7B, the battery 78 discharges. Therefore, as shown in FIG. 7D, the SOC decreases. In addition, before the aircraft takes off, the aircraft is not disturbed at the time point. Therefore, the electric power request value does not change. In addition, the electric energy required for starting GT72, that is, the reduction amount of the SOC, can be estimated. Therefore, until the aircraft takes off, the target SOC is equal to the SOC. Therefore, the target SOC follows the SOC.

[0069] During the period from time point t3 to time point t4, the flight mode is the vertical takeoff mode M1. During the vertical takeoff mode M1, the controller 64 controls the PCU 74 and the driver 66 to supply the generated power of the MG 76 to the electric motor 68. During this period, as shown in FIG. 7B, the battery 78 does not charge and discharge stably. However, the battery 78 is charged with the excess part of the generated power of the MG 76 exceeding the electric power request value and is discharged according to the shortage part of the generated power of the MG 76 relative to the electric power request value. The electric power change value during the vertical takeoff mode M1 is large. Therefore, the excess or shortage amount of the generated power of the MG 76 relative to the electric power request value is also large. As shown in FIG. 7D, the controller 64 (SOC calculator 104) sets the lowest value T1 among the target values set during flight as the target SOC. Accordingly, the battery 78 can be charged with the excess generated power. Therefore, overcharging of the battery 78 is suppressed.

[0070] During the time period from time point t4 to time point t5, the flight mode is the ascending mode M2. During the ascending mode M2, the controller 64 controls the PCU 74 and the driver 66 to supply the electric power generated by the MG 76 to the electric motor 68. During this period, as shown in FIG. 7B, the battery 78 does not stably charge and discharge. On the other hand, the battery 78 is charged with the excess of the electric power generated by the MG 76 over the electric power request value, or discharged according to the shortage of the electric power generated by the MG 76 relative to the electric power request value. The variation value of the electric power during the ascending mode M2 is the second largest. Therefore, the excess or shortage of the electric power generated by the MG 76 relative to the electric power request value is also the second largest. As shown in FIG. 7D, the controller 64 sets the second lowest value T2 among the target values set during flight as the target SOC. Accordingly, the battery 78 can be charged with the excess of the generated power. Therefore, overcharging of the battery 78 is suppressed.

[0071] During the time period from time point t5 to time point t6, the flight mode is the cruising mode M3. During the cruising mode M3, the controller 64 controls the PCU 74 and the driver 66 to supply the electric power generated by the MG 76 to the electric motor 68. During this period, the power consumption of the electric motor 68 is small. Therefore, as shown in FIG. 7B, the battery 78 stably charges. However, at the time point when the battery 78 is fully charged, the charging of the battery 78 stops. The variation value of the electric power during the cruising mode M3 is small. Therefore, the excess or shortage of the electric power generated by the MG 76 relative to the electric power request value is also small. As shown in FIG. 7D, the controller 64 sets the second highest value T3 among the target values set during flight as the target SOC. Accordingly, the SOC of the battery 78 can be increased. Therefore, the available time of the battery 78 can be extended.

[0072] During the time period from time point t6 to time point t7, the flight mode is the descending mode M4. During the descending mode M4, the controller 64 controls the PCU 74 and the driver 66 to supply the electric power generated by the MG 76 to the electric motor 68. During this period, the power consumption of the electric motor 68 is small. As shown in FIG. 7B, the battery 78 does not stably charge and discharge. The variation value of the electric power during the descending mode M4 is small. Therefore, the excess or shortage of the electric power generated by the MG 76 relative to the electric power request value is also small. As shown in FIG. 7D, the controller 64 sets the highest value T4 among the target values set during flight as the target SOC. Accordingly, the SOC of the battery 78 can be increased.

[0073] During the time period from time point t7 to time point t8, the flight mode is the vertical landing mode M5. During the vertical landing mode M5, the controller 64 controls the PCU 74 and the driver 66 to supply the generated power of the MG 76 to the electric motor 68. During this period, as shown in FIG. 7B, the battery 78 does not stably charge and discharge. However, the battery 78 is charged with the excess of the generated power of the MG 76 over the electric power request value and discharged according to the shortage of the generated power of the MG 76 relative to the electric power request value. The variation value of the electric power during the vertical landing mode M5 is large. Therefore, the excess or shortage of the generated power of the MG 76 relative to the electric power request value is also large. As shown in FIG. 7D, the controller 64 sets the lowest value T1 among the target values set during flight as the target SOC. Accordingly, the battery 78 can be charged with the excess of the generated power. Therefore, overcharging of the battery 78 is suppressed.

[0074] [4.2 Second and Third Control Examples]

[0075] The second control example and the third control example are described with reference to FIGS. 8A to 8D. Compared with the first control example, the control before takeoff is different in the second control example. Compared with the first control example, the control after landing is different in the third control example.

[0076] In the first control example, during the time period from time point t2 to time point t3, the controller 64 controls the PCU 74 and the driver 66 to consume the electric power of the battery 78. Therefore, the battery 78 discharges and the SOC decreases.

[0077] In the second control example, during the time period from time point t2 to time point t3, the controller 64 controls the PCU 74 and the driver 66 to supply the generated power of the MG 76 to the battery 78. Before the flight mode transfers to the vertical takeoff mode M1, the controller 64 sets a value larger than the SOC after the decrease as the target SOC. For example, the controller 64 sets the set value T1 of the vertical takeoff mode M1 as the target SOC. In this way, the SOC that has decreased due to the start of the GT 72 is restored earlier. Therefore, after takeoff, the SOC can quickly approach the target SOC.

[0078] In the third control example, at time point t81 between time point t8 and time point t9, the controller 64 controls the PCU 74 and the driver 66 to supply the generated power of the MG 76 to the battery 78. In this control, the controller 64 sets a value larger than the set value T1 of the vertical takeoff mode M1 as the target SOC. In this way, the decrease in the SOC of the battery 78 due to the start of the GT 72 can be replenished in advance. Therefore, after takeoff, the SOC can quickly approach the target SOC.

[0079] [4.3 Other Control Examples]

[0080] In the first to third control examples, the storage unit 94 of the flight controller 80 stores in advance the relationships among the flight mode, the electric power request value, and the electric power change value. The storage unit 94 may also store in advance the relationships between the measured values of each of the wind speed, wind direction, altitude, temperature, and atmospheric pressure and the electric power request value. The storage unit 94 may also store in advance the relationships between the measured values of each of the wind speed, wind direction, altitude, temperature, and atmospheric pressure and the electric power change value. The arithmetic unit 92 of the flight controller 80 may calculate the electric power request value based on the measured values of each of the wind speed, wind direction, altitude, temperature, and atmospheric pressure. The arithmetic unit 92 of the flight controller 80 may calculate the electric power change value based on the measured values of each of the wind speed, wind direction, altitude, temperature, and atmospheric pressure.

[0081] When the steady wind speed increases, the amount of change in the attitude of the aircraft body and the electric power request value are greater than those in the no-wind state. In addition, when the amount of change in the wind speed increases, the amount of change in the attitude of the aircraft body caused by the fluctuating wind increases. Therefore, the electric power required for attitude recovery increases. That is, the electric power request value increases. In addition, there is generally a correlation between the steady wind and the fluctuating wind. When the steady wind speed is small, the amount of change in the wind speed is also small, and when the steady wind speed is large, the amount of change is also large. That is, when the steady wind speed is weak, the electric power request value is small, the fluctuating wind speed is also small, and therefore, the electric power change value is also small. On the other hand, when the steady wind speed is large, the electric power request value is large, the fluctuating wind speed is also large, and therefore, the electric power change value is also large. In this way, the electric power request value and the electric power change value change according to the wind speed. By calculating the electric power request value and the electric power change value corresponding to the wind speed by the arithmetic unit 92, it is possible to cope with the electric power change when fluctuating wind occurs.

[0082] In addition, when the altitude is high, the fluctuating wind becomes stronger. Therefore, the storage unit 94 may also store in advance the electric power request value and the electric power change value corresponding to the altitude. In addition, the storage unit 94 may also store in advance the electric power request value corresponding to the air density determined based on at least one of the altitude, temperature, and atmospheric pressure. The storage unit 94 may also store in advance the electric power change value corresponding to the air density determined based on at least one of the altitude, temperature, and atmospheric pressure.

[0083] [5 Invention obtained according to the embodiment]

[0084] The invention that can be grasped according to the above embodiment is described below.

[0085] An aspect of the present invention is a hybrid aircraft 10 that uses a plurality of rotors (takeoff and landing rotor 20, cruise rotor 22) for flight,

[0086] and includes a generator (MG76), a battery 78, a plurality of electric motors 68, and a controller 64, wherein,

[0087] the battery 78 is charged with the electric power generated by the generator;

[0088] A plurality of the electric motors 68 rotate respective rotors by electric power generated by the generator and electric power supplied from the battery 78.

[0089] The controller 64 sets a target remaining capacity (target SOC) of the battery 78 according to the flight state of the aircraft body, and controls charging and discharging of the battery 78 so that the remaining capacity (SOC) of the battery 78 approaches the target remaining capacity.

[0090] In the above structure, the target SOC is set according to the flight state. According to the above structure, in a flight state where the excess or deficiency amount of the electric power request value with respect to the generated power is large, overcharging of the battery 78 can be prevented by setting the target SOC to be low. Also, according to the above structure, charging can be performed with the generated power without waste. Further, according to the above structure, in a flight state where the excess or deficiency amount of the electric power request value with respect to the generated power is small, insufficient charging of the battery 78 can be prevented by setting the target SOC to be high.

[0091] In an aspect of the present invention, it may also be that

[0092] The controller 64 estimates an electric power change value according to the flight state of the aircraft body, and determines the target remaining capacity according to the electric power change value, where the electric power change value is a change value of the electric power used by a plurality of the electric motors 68.

[0093] In an aspect of the present invention, it may also be that

[0094] The controller 64 determines a flight mode as the flight state,

[0095] The flight mode includes a vertical takeoff and landing mode (M1, M5) and a horizontal flight mode (M3), where the vertical takeoff and landing mode (M1, M5) is a mode in which a part of the rotors generate lift to move the aircraft body in the vertical direction; the horizontal flight mode (M3) is a mode in which a part of the rotors generate thrust to move the aircraft body in the horizontal direction.

[0096] According to the above structure, the target SOC is set according to the flight mode related to the electric power change value. Therefore, overcharging of the battery 78 can be prevented. Insufficient charging of the battery 78 can be prevented.

[0097] In an aspect of the present invention, it may also be that

[0098] There is a storage unit 98 that stores a table (SOC conversion table TB) establishing a correspondence between the flight state and the electric power change value.

[0099] The controller 64 uses the table to estimate the value of the change in electric power.

[0100] In an aspect of the present invention, it may also be that

[0101] Before the aircraft takes off, the controller 64 operates the generator to charge the battery 78, and during the period from when the aircraft lands until it takes off, replenishes the decrease in the remaining capacity when the battery 78 is used.

[0102] According to the above structure, the SOC of the battery 78 is relatively high at takeoff. Therefore, after takeoff, the SOC can quickly approach the target SOC.

[0103] In an aspect of the present invention, it may also be that

[0104] It has an internal combustion engine (GT72) that supplies power to the generator.

[0105] The generator is a motor generator 76. When the internal combustion engine starts, the motor generator 76 operates the internal combustion engine by receiving the electric power supplied from the battery 78, and after the internal combustion engine starts, generates electricity by the power generated by the internal combustion engine.

[0106] During the period from when the internal combustion engine starts until the aircraft takes off, the controller 64 operates the generator to charge the battery 78, and replenishes the decrease in the remaining capacity caused by the start of the internal combustion engine.

[0107] According to the above structure, it is possible to replenish the decrease in SOC caused by the start of GT72 before takeoff.

[0108] In an aspect of the present invention, it may also be that

[0109] It has an internal combustion engine (GT72) that supplies power to the generator.

[0110] The generator is a motor generator 76. When the internal combustion engine starts, the motor generator 76 operates by receiving the electric power supplied from the battery 78, and after the internal combustion engine starts, generates electricity by the power generated by the internal combustion engine.

[0111] During the period from when the aircraft lands until the internal combustion engine stops, the controller 64 operates the generator to pre-charge the battery 78 in advance, and pre-replenishes the decrease in the remaining capacity caused by the start of the internal combustion engine.

[0112] According to the above structure, it is possible to supplement the reduction in SOC caused by the start of GT72 before takeoff.

[0113] In the embodiment of the present invention, it may also be that

[0114] the controller 64 sets the target remaining capacity to be lower when the aircraft moves in the vertical direction, and sets the target remaining capacity to be higher when the aircraft moves in the horizontal direction.

[0115] In addition, the hybrid aircraft according to the present invention is not limited to the foregoing embodiments, and various structures can be adopted without departing from the gist of the present invention.

Claims

1. A hybrid aircraft that uses multiple rotors for flight, characterized in that it has a generator, a battery, multiple electric motors, and a controller, wherein the battery is charged by the electric power generated by the generator; the multiple electric motors rotate each of the rotors by the electric power generated by the generator and the electric power supplied by the battery; the controller estimates the electric power variation value according to the flight state of the aircraft, sets the target remaining capacity of the battery based on the electric power variation value, and controls the charge and discharge of the battery so that the remaining capacity of the battery approaches the target remaining capacity, where the electric power variation value is the variation value of the electric power used by the multiple electric motors.

2. The hybrid aircraft according to claim 1, characterized in that the controller determines the flight mode as the flight state, the flight mode includes a vertical takeoff and landing mode and a horizontal flight mode, wherein the vertical takeoff and landing mode is a mode in which lift is generated by a part of the rotors to move the aircraft in the vertical direction; the horizontal flight mode is a mode in which thrust is generated by a part of the rotors to move the aircraft in the horizontal direction.

3. The hybrid aircraft according to claim 1, characterized in that it has a storage unit that stores a table establishing a correspondence between the flight state and the electric power variation value, the controller uses the table to estimate the electric power variation value.

4. The hybrid aircraft according to claim 1, characterized in that the controller operates the generator to charge the battery before the aircraft takes off, and supplements the reduction amount of the remaining capacity when the battery is used during the period from the aircraft landing to taking off.

5. The hybrid aircraft according to claim 4, characterized in that it has an internal combustion engine that supplies power to the generator, the generator is an electric generator that operates the internal combustion engine by receiving the electric power supplied by the battery when the internal combustion engine starts, and generates electricity by the power generated by the internal combustion engine after the internal combustion engine starts, the controller operates the generator to charge the battery during the period from the start of the internal combustion engine to the takeoff of the aircraft, and supplements the reduction amount of the remaining capacity accompanying the start of the internal combustion engine.

6. The hybrid aircraft according to claim 4, characterized in that it has an internal combustion engine that supplies power to the generator, the generator is an electric generator that operates the internal combustion engine by receiving the electric power supplied by the battery when the internal combustion engine starts, and generates electricity by the power generated by the internal combustion engine after the internal combustion engine starts, the controller operates the generator to pre-charge the battery during the period from the landing of the aircraft to the stop of the internal combustion engine, and pre-supplements the reduction amount of the remaining capacity accompanying the start of the internal combustion engine.

7. The hybrid aircraft according to any one of claims 1 to 6, characterized in that The controller sets the target remaining capacity to be relatively low when the body moves in the vertical direction, and sets the target remaining capacity to be relatively high when the body moves in the horizontal direction.

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