CONTROL DEVICE FOR AIRCRAFT

When a first generator is stopped, an electric motor control section reduces the output power of a VTOL electric motor and a cruise electric motor that operate with electric power supplied from the first generator, and increases the output power of a VTOL electric motor and a cruise electric motor that operate with electric power supplied from a second generator.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is based upon and claims the benefit of priority from Japanese Patent Application No. 2022-054782 filed on Mar. 30, 2022, the contents of which are incorporated herein by reference.

BACKGROUND OF THE INVENTION

Field of the Invention

The present invention relates to a control device for an aircraft.

Description of the Related Art

US 2021/0370786 A1 discloses a multi-rotor aircraft. The multi-rotor aircraft includes a plurality of rotors, and an electric motor is provided corresponding to each rotor. Some of the electric motors are supplied with electric power from one of two generators, and another some of the electric motors are supplied with electric power from the other generator.

SUMMARY OF THE INVENTION

In the technique disclosed in US 2021/0370786 A1, when electric power cannot be supplied from one of the generators to some of the electric motors, these electric motors are supplied with electric power from a battery. In order to increase the flight duration of the multi-rotor aircraft after electric power is no longer supplied from the generator, the capacity of the battery needs to be increased. However, when the capacity of the battery is increased, there arises a problem in that the weight of the battery increases and the weight of the fuselage also increases.

An object of the present invention is to solve the above-mentioned problem.

According to an aspect of the present invention, provided is a control device for an aircraft, the aircraft comprising: at least one first generator configured to generate electric power; at least one first battery configured to store electric power; at least one first electric motor configured to operate with electric power supplied from the first generator and the first battery; at least one second generator configured to generate electric power; at least one second battery configured to store electric power; at least one second electric motor configured to operate with electric power supplied from the second generator and the second battery; and a plurality of rotors configured to generate thrust acting on a fuselage, wherein the control device comprises an electric motor control section configured to control the first electric motor and the second electric motor, and wherein each of the rotors is driven by one of the first electric motor or the second electric motor, or both of the first electric motor and the second electric motor, and in a case where supply of electric power from the first generator to the first electric motor is unavailable and supply of electric power from the second generator to the second electric motor is available, the electric motor control section reduces power consumption of the first electric motor by reducing thrust generated by the rotor driven by the first electric motor, and increases thrust generated by the rotor driven by the second electric motor, as compared with a case where the supply of the electric power from the first generator to the first electric motor is available and the supply of the electric power from the second generator to the second electric motor is available.

According to the present invention, an increase in the weight of the fuselage can be suppressed.

DETAILED DESCRIPTION OF THE INVENTION

First Embodiment

Configuration of Aircraft

FIG.1is a schematic diagram of an aircraft10. The aircraft10of the present embodiment is an electric vertical take-off and landing aircraft (eVTOL aircraft). In the aircraft10, rotors are driven by electric motors. The aircraft10generates vertical thrust and horizontal thrust by the rotors. Further, the aircraft10is a hybrid aircraft. The aircraft10includes a generator and a battery as power sources of the electric motor. In the aircraft10, electric power generated by the generator is supplied to the electric motor. When the electric power generated by the generator is insufficient with respect to the required electric power, the electric power stored in the battery is supplied to the electric motor.

The aircraft10includes a fuselage12. The fuselage12is provided with a cockpit, a cabin, and the like. A pilot rides in the cockpit and controls the aircraft10. Passengers and the like ride in the cabin. The aircraft10may be automatically controlled.

The aircraft10includes a front wing14and a rear wing16. When the aircraft10moves forward, lift is generated in each of the front wing14and the rear wing16.

The aircraft10includes eight VTOL rotors18V. The eight VTOL rotors18V are a rotor18V1, a rotor18V2, a rotor18V3, a rotor18V4, a rotor18V5, a rotor18V6, a rotor18V7, and a rotor18V8.

The rotor18V1, the rotor18V3, the rotor18V5, and the rotor18V7are disposed on the left side of a center line A of the fuselage12in the left-right direction. The rotor18V2, the rotor18V4, the rotor18V6, and the rotor18V8are disposed on the right side of the center line A. That is, four VTOL rotors18V are disposed on the left side of the center line A, and four VTOL rotors18V are disposed on the right side of the center line A.

The center of gravity G of the aircraft10is located on the center line A of the fuselage12. In a state where the aircraft10is viewed from above, the center of gravity G is located between the rotor18V4and the rotor18V6in the front-rear direction of the fuselage12. Further, the center of gravity G is located between the rotor18V3and the rotor18V5in the front-rear direction of the fuselage12.

In a state where the aircraft10is viewed from above, the rotor18V8is provided at a position point-symmetrical to the rotor18V1with respect to the center of gravity G. The rotor18V7is provided at a position point-symmetrical to the rotor18V2with respect to the center of gravity G. The rotor18V6is provided at a position point-symmetrical to the rotor18V3with respect to the center of gravity G. The rotor18V5is provided at a position point-symmetrical to the rotor18V4with respect to the center of gravity G.

One VTOL electric motor20V is provided for one VTOL rotor18V. Specifically, an electric motor20V1_1is provided for the rotor18V1, an electric motor20V2_2is provided for the rotor18V2, an electric motor20V3_2is provided for the rotor18V3, an electric motor20V4_1is provided for the rotor18V4, an electric motor20V5_1is provided for the rotor18V5, an electric motor20V6_2is provided for the rotor18V6, an electric motor20V7_2is provided for the rotor18V7, and an electric motor20V8_1is provided for the rotor18V8. Each VTOL rotor18V is driven by each VTOL electric motor20V.

Each VTOL rotor18V generates thrust mainly toward the upper side of the fuselage12. The thrust of each VTOL rotor18V is controlled by adjusting the rotational speed of the rotor and the pitch angle of the blades. If the conditions such as the pitch angle of the blades are constant, the thrust generated by the VTOL rotor18V increases as the output power of the VTOL electric motor20V increases. Each VTOL rotor18V is mainly used during vertical take-off, during transition from vertical take-off to cruising, during transition from cruising to vertical landing, during vertical landing, during hovering, and the like. Further, each VTOL rotor18V is used during attitude control.

By controlling the thrust of each VTOL rotor18V, a propulsive force is applied mainly upward to the fuselage12. By controlling the thrust of each VTOL rotor18V, a roll moment, a pitch moment, and a yaw moment are caused to act on the fuselage12. The VTOL rotor18V corresponds to a vertical rotor of the present invention.

The aircraft10includes two cruise rotors22C. The two cruise rotors22C are a rotor22C1and a rotor22C2. The rotor22C1and the rotor22C2are attached to the rear portion of the fuselage12. The rotor22C1is disposed on the left side of the center line A. The rotor22C2is disposed on the right side of the center line A. That is, one cruise rotor22C is disposed on the left side of the center line A, and one cruise rotor22C is disposed on the right side of the center line A.

Two cruise electric motors24C are provided for each cruise rotor22C. Specifically, an electric motor24C1_1and an electric motor24C2_2are provided for the rotor22C1, and an electric motor24C3_1and an electric motor24C4_2are provided for the rotor22C2. One cruise rotor22C is driven by two cruise electric motors24C.

Each cruise rotor22C generates thrust mainly toward the front of the fuselage12. The thrust of each cruise rotor22C is controlled by adjusting the rotational speed of the rotor and the pitch angle of the blades. If the conditions such as the pitch angle of the blades are constant, the thrust generated by the cruise rotor22C increases as the output power of the cruise electric motor24C increases. Each cruise rotor22C is mainly used during transition from vertical take-off to cruising, during cruising, during transition from cruising to vertical landing, and the like. By controlling the thrust of each cruise rotor22C, a propulsive force is applied mainly forward to the fuselage12. The cruise rotor22C corresponds to a horizontal rotor of the present invention.

Configuration of Power Supply System

FIG.2is a schematic diagram showing a configuration of a power supply system26.

The aircraft10includes the electric motor20V1_1, the electric motor20V4_1, the electric motor20V5_1, the electric motor20V8_1, the electric motor24C1_1, and the electric motor24C3_1as drive sources of a first drive system28. Hereinafter, the electric motor20V1_1, the electric motor20V4_1, the electric motor20V5_1, and the electric motor20V8_1may be referred to as a VTOL electric motor20V of the first drive system28. Further, the electric motor24C1_1and the electric motor24C3_1may be referred to as a cruise electric motor24C of the first drive system28. The VTOL electric motor20V of the first drive system28and the cruise electric motor24C of the first drive system28correspond to a first electric motor of the present invention.

The aircraft10includes the electric motor20V2_2, the electric motor20V3_2, the electric motor20V6_2, the electric motor20V7_2, the electric motor24C2_2, and the electric motor24C4_2as drive sources of a second drive system30. Hereinafter, the electric motor20V2_2, the electric motor20V3_2, the electric motor20V6_2, and the electric motor20V7_2may be referred to as a VTOL electric motor20V of the second drive system30. Further, the electric motor24C2_2and the electric motor24C4_2may be referred to as a cruise electric motor24C of the second drive system30. The VTOL electric motor20V of the second drive system30and the cruise electric motor24C of the second drive system30correspond to a second electric motor of the present invention.

The power supply system26includes two main power source devices32and four auxiliary power source devices34. The power supply system26supplies electric power to four load modules36.

The two main power source devices32include a first main power source device32aand a second main power source device32b.The four auxiliary power source devices34include a first auxiliary power source device34a,a second auxiliary power source device34b,a third auxiliary power source device34c, and a fourth auxiliary power source device34d.The four load modules36include a first load module36a,a second load module36b,a third load module36c,and a fourth load module36d.

The power supply system26includes two power supply circuits38. The two power supply circuits38include a first power supply circuit38aand a second power supply circuit38b. The first power supply circuit38aand the second power supply circuit38bare not connected to each other and are provided independently.

Each power supply circuit38includes a main power source circuit40and an auxiliary power source circuit42. The main power source circuit40is provided for each main power source device32. The auxiliary power source circuit42is provided for each auxiliary power source device34.

Each main power source device32includes a gas turbine44, a generator46, and a power control unit (hereinafter referred to as PCU)48. The gas turbine44drives the generator46. As a result, the generator46generates electric power. The PCU48converts the AC power generated by the generator46into DC power and outputs the DC power to the main power source circuit40. When starting the gas turbine44, the PCU48converts the DC power supplied by the main power source circuit40into AC power, and outputs the AC power to the generator46. The generator46is operated by the AC power input from the PCU48, and the generator46drives the gas turbine44.

Hereinafter, the generator46in the first main power source device32amay be referred to as a first generator46a. Further, the generator46in the second main power source device32bmay be referred to as a second generator46b.

Each auxiliary power source device34includes a battery50. The battery50is charged with DC power supplied from the main power source device32. Hereinafter, the battery50in the first auxiliary power source device34amay be referred to as a first battery50a.Further, the battery50in the second auxiliary power source device34bmay be referred to as a second battery50b,the battery50in the third auxiliary power source device34cmay be referred to as a third battery50c,and the battery50in the fourth auxiliary power source device34dmay be referred to as a fourth battery50d.

The first battery50aand the second battery50bsupply electric power to the VTOL electric motors20V of the first drive system28and the cruise electric motors24C of the first drive system28. That is, the first battery50aand the second battery50bfunction as power storage devices of the first drive system28. Hereinafter, each of the first battery50aand the second battery50bmay be referred to as a battery50of the first drive system28. The third battery50cand the fourth battery50dsupply electric power to the VTOL electric motors20V of the second drive system30and the cruise electric motors24C of the second drive system30. That is, the third battery50cand the fourth battery50dfunction as power storage devices of the second drive system30. Hereinafter, each of the third battery50cand the fourth battery50dmay be referred to as a battery50of the second drive system30. The battery50of the first drive system28corresponds to a first battery of the present invention. The battery50of the second drive system30corresponds to a second battery of the present invention.

Each load module36includes two VTOL drive units52and one cruise drive unit54.

Each VTOL drive unit52includes an inverter56and the VTOL electric motor20V. The inverter56converts the DC power supplied from the main power source circuit40into three phase AC power, and outputs the three phase AC power to the VTOL electric motor20V.

The cruise drive unit54includes an inverter58and the cruise electric motor24C. The inverter58converts the DC power supplied from the main power source circuit40into three phase AC power, and outputs the three phase AC power to the cruise electric motor24C.

Each of the first load module36aand the third load module36cincludes a converter60. The converter60steps down the voltage of the DC power supplied from the main power source device32, and outputs the stepped-down power to a device operated by DC power. The device operated by DC power is, for example, a cooling device that cools the PCU48, the inverters56and58, and the like.

Each main power source circuit40includes one common bus62, one contactor unit64, two contactor units66, one current sensor68, and two current sensors70.

The common bus62connects one main power source device32and two load modules36. The common bus62connects the two load modules36in parallel with the main power source device32.

The contactor unit64is provided between the main power source device32and the common bus62. The contactor unit64switches between a conduction state in which current flows between the main power source device32and the common bus62, and an interruption state in which the flow of the current between the main power source device32and the common bus62is interrupted. The contactor unit64includes a contactor64aand a contactor64b.The contactor64ais provided on a positive line of the main power source circuit40. The contactor64bis provided on a negative line of the main power source circuit40. The contactor unit64may include only one of the contactor64aor the contactor64b.

Each contactor unit66is provided between each load module36and the common bus62. The contactor unit66switches between a conduction state in which current flows between each load module36and the common bus62, and an interruption state in which the flow of the current between each load module36and the common bus62is interrupted. The contactor unit66includes a contactor66aand a contactor66b.The contactor66ais provided on the positive line of the main power source circuit40. The contactor66bis provided on the negative line of the main power source circuit40. The contactor unit66may include only one of the contactor66aor the contactor66b.When the contactor unit64includes only the contactor64a,each contactor unit66preferably includes only the contactor66b. When the contactor unit64includes only the contactor64b,each contactor unit66preferably includes only the contactor66a.

The current sensor68is provided between the contactor unit64and the common bus62. The current sensor68is provided on the positive line of the main power source circuit40. Each current sensor70is provided between each contactor unit66and the common bus62. Each current sensor70is provided on the positive line of the main power source circuit40.

Each auxiliary power source circuit42is connected to both the main power source circuit40and the load module36. The auxiliary power source circuit42supplies electric power from the auxiliary power source device34to the load module36. The auxiliary power source circuit42includes a contactor unit72and a current sensor74.

The contactor unit72is provided between the auxiliary power source device34and the load module36. The contactor unit72switches between a conduction state in which current flows between the auxiliary power source device34and the load module36, and an interruption state in which the flow of the current between the auxiliary power source device34and the load module36is interrupted. The contactor unit72includes a contactor72a,a contactor72b,and a precharge circuit72c.The contactor72ais provided on a positive line of the auxiliary power source circuit42. The contactor72bis provided on a negative line of the auxiliary power source circuit42. The precharge circuit72cis provided in parallel with the contactor72b.The precharge circuit72cincludes a contactor72dand a resistor72e.The current sensor74is provided on the negative line of the auxiliary power source circuit42.

The contactor unit72may include only the contactor72band the precharge circuit72c.The precharge circuit72cmay be provided in parallel with the contactor72a.In this case, the contactor unit72may include only the contactor72aand the precharge circuit72c.

A diode76is provided between the main power source circuit40and each auxiliary power source circuit42. An anode of the diode76is connected to the main power source circuit40, and a cathode of the diode76is connected to the auxiliary power source circuit42. The diode76allows electric power to be supplied from the main power source circuit40to the auxiliary power source circuit42. The diode76prevents electric power from being supplied from the auxiliary power source circuit42to the main power source circuit40. When the main power source circuit40is short-circuited, electricity is prevented from flowing from the auxiliary power source device34to the main power source circuit40. As a result, even when the main power source circuit40is short-circuited, electric power can be supplied from the auxiliary power source device34to the load module36.

A transistor78is provided in parallel with the diode76. When the transistor78is ON, electric power is supplied from the auxiliary power source device34to the main power source circuit40while bypassing the diode76. The generator46is operated by the electric power supplied from the auxiliary power source device34, and the gas turbine44can be started.

Among the VTOL rotors18V disposed on the left side of the center line A of the fuselage12(FIG.1), the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28are two rotors18V1and18V5(FIG.2). Among the VTOL rotors18V disposed on the left side of the center line A of the fuselage12(FIG.1), the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30are two rotors18V3and18V7(FIG.2). That is, among the VTOL rotors18V disposed on the left side of the center line A of the fuselage12, the number of the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28is equal to the number of the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30.

Among the VTOL rotors18V disposed on the right side of the center line A of the fuselage12(FIG.1), the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28are two rotors18V4and18V8(FIG.2). Among the VTOL rotors18V disposed on the right side of the center line A of the fuselage12(FIG.1), the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30are two rotors18V2and18V6(FIG.2). That is, among the VTOL rotors18V disposed on the right side of the center line A of the fuselage12, the number of the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28is equal to the number of the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30.

The rotor22C1disposed on the left side of the center line A of the fuselage12is driven by the cruise electric motor24C of the first drive system28and is driven by the cruise electric motor24C of the second drive system30(FIG.2). That is, regarding the cruise rotors22C disposed on the left side of the center line A of the fuselage12, the number of the cruise rotors22C driven by the cruise electric motor24C of the first drive system28is equal to the number of the cruise rotors22C driven by the cruise electric motor24C of the second drive system30.

The rotor22C2disposed on the right side of the center line A of the fuselage12is driven by the cruise electric motor24C of the first drive system28and is driven by the cruise electric motor24C of the second drive system30(FIG.2). That is, regarding the cruise rotors22C disposed on the right side of the center line A of the fuselage12, the number of the cruise rotors22C driven by the cruise electric motor24C of the first drive system28is equal to the number of the cruise rotors22C driven by the cruise electric motor24C of the second drive system30.

Configuration of Flight Controller

The power supply system26includes a flight controller80. The flight controller80controls the thrust output from each VTOL rotor18V and each cruise rotor22C.FIG.3is a control block diagram of the flight controller80.

The flight controller80includes a computation section82and a storage section84. The computation section82is, for example, a processor such as a central processing unit (CPU) or a graphics processing unit (GPU). The computation section82includes an output power command value calculation section86, a main power source device monitoring section88, and an electric motor control section90. The output power command value calculation section86, the main power source device monitoring section88, and the electric motor control section90are realized by the computation section82executing programs stored in the storage section84. At least part of the output power command value calculation section86, the main power source device monitoring section88, and the electric motor control section90may be realized by an integrated circuit such as an application specific integrated circuit (ASIC) or a field-programmable gate array (FPGA). At least part of the output power command value calculation section86, the main power source device monitoring section88, and the electric motor control section90may be realized by an electronic circuit including a discrete device.

The storage section84is configured by a volatile memory (not shown) and a non-volatile memory (not shown) which are computer-readable storage media. The volatile memory is, for example, a random access memory (RAM) or the like. The non-volatile memory is, for example, a read only memory (ROM), a flash memory, or the like. Data and the like are stored in, for example, the volatile memory. Programs, tables, maps, and the like are stored in, for example, the non-volatile memory. At least a part of the storage section84may be included in the processor, the integrated circuit, or the like described above.

The output power command value calculation section86calculates an output power command value for each VTOL electric motor20V, and an output power command value for each cruise electric motor24C. The output power command value is determined in accordance with an operation amount of an operation input section by the pilot. The operation input section is, for example, a control stick, a pedal, a lever, or the like. The operation amount of the operation input section and the output power command value may not have a one to-one correspondence. The output power command value may be variable with respect to the operation amount of the operation input section in accordance with the operation range of the operation input section, the operation speed of the operation input section, the attitude of the fuselage12, and the like.

When there is no operation input to the operation input section by the pilot, the output power command value may be automatically determined and hovering may be performed regardless of the operation amount of the operation input section. Further, when the aircraft10is automatically controlled, the output power command value may be automatically determined in accordance with a preset flight path, regardless of the operation amount of the operation input section.

The main power source device monitoring section88monitors the state of each main power source device32. For example, the main power source device monitoring section88monitors whether or not the generator46is operating, as the state of the main power source device32. When the generator46is stopped, electric power cannot be supplied from the generator46to each VTOL electric motor20V and each cruise electric motor24C.

The electric motor control section90controls each VTOL electric motor20V to set the output power of each VTOL electric motor20V to the power command value. The electric motor control section90controls each cruise electric motor24C to set the output power of each cruise electric motor24C to the power command value. When a predetermined condition is satisfied, the electric motor control section90changes the output power distribution in each VTOL electric motor20V and each cruise electric motor24C.

Output Power Control

FIG.4is a flowchart showing a process of output power control. The output power control is performed by the electric motor control section90. The output power control is repeatedly executed at a predetermined cycle while the aircraft10is activated.

In step S1, the electric motor control section90determines whether or not both the first generator46aand the second generator46bare operating. When both the first generator46aand the second generator46bare operating, the process proceeds to step S2. When one of the first generator46aor the second generator46bis stopped, the process proceeds to step S3.

In step S2, the electric motor control section90controls each VTOL electric motor20V and each cruise electric motor24C based on the output power command value. Thereafter, the output power control is ended. By the process of step S2, the output power of each VTOL electric motor20V and each cruise electric motor24C becomes substantially equal to the power command value.

In step S3, the electric motor control section90determines whether the stopped generator is the first generator46aor the second generator46b.When the first generator46ais stopped, the process proceeds to step S4. When the second generator46bis stopped, the process proceeds to step S8.

When the first generator46ais stopped, the electric motor control section90changes the output power distribution in steps S4to S7.

In step S4, the electric motor control section90makes the output power of each VTOL electric motor20V of the first drive system28smaller than the output power command value. Thereafter, the process proceeds to step S5. By the process of step S4, the power consumption of each VTOL electric motor20V of the first drive system28decreases. In addition, if other conditions are constant, the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28decreases.

In step S5, the electric motor control section90makes the output power of each VTOL electric motor20V of the second drive system30larger than the output power command value. Thereafter, the process proceeds to step S6. By the process of step S5, the power consumption of each VTOL electric motor20V of the second drive system30increases. In addition, if the other conditions are constant, the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30increases.

In step S5, the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30is increased by an amount corresponding to the decrease in the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28in step S4. As a result, the sum of the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28and the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30is maintained before and after the change of the output power distribution.

In step S6, the electric motor control section90sets the output power of each cruise electric motor24C of the first drive system28to 0. Thereafter, the process proceeds to step S7. By the process of step S6, the power consumption of each cruise electric motor24C of the first drive system28decreases. The thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the first drive system28becomes 0.

In step S7, the electric motor control section90makes the output power of each cruise electric motor24C of the second drive system30larger than the output power command value. Thereafter, the output power control is ended. By the process of step S7, the power consumption of each cruise electric motor24C of the second drive system30increases. In addition, if the other conditions are constant, the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the second drive system30increases. The sum of the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the first drive system28and the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the second drive system30is decreased as compared with that before the change of the output power distribution.

When the second generator46bis stopped, the electric motor control section90changes the output power distribution in steps S8to S11.

In step S8, the electric motor control section90makes the output power of each VTOL electric motor20V of the second drive system30smaller than the output power command value. Thereafter, the process proceeds to step S9. By the process of step S8, the power consumption of each VTOL electric motor20V of the second drive system30decreases. In addition, if other conditions are constant, the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30decreases.

In step S9, the electric motor control section90makes the output power of each VTOL electric motor20V of the first drive system28larger than the output power command value. Thereafter, the process proceeds to step S10. By the process of step S9, the power consumption of each VTOL electric motor20V of the first drive system28increases. In addition, if the other conditions are constant, the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28increases.

In step S9, the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28is increased by an amount corresponding to the decrease in the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30in step S8. As a result, the sum of the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28and the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30is maintained before and after the change of the output power distribution.

In step S10, the electric motor control section90sets the output power of each cruise electric motor24C of the second drive system30to 0. Thereafter, the process proceeds to step S11. By the process of step S10, the power consumption of each cruise electric motor24C of the second drive system30decreases. The thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the second drive system30becomes 0.

In step S11, the electric motor control section90makes the output power of each cruise electric motor24C of the first drive system28larger than the output power command value. Thereafter, the output power control is ended. By the process of step S11, the power consumption of each cruise electric motor24C of the first drive system28increases. In addition, if the other conditions are constant, the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the first drive system28increases. The sum of the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the first drive system28and the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the second drive system30is decreased as compared with that before the change of the output power distribution.

FIGS.5and6are image diagrams of output power distribution in each VTOL electric motor20V and each cruise electric motor24C. The numerical values of percentages shown inFIGS.5and6indicate the ratio of the output power to the output power command value for each VTOL electric motor20V and each cruise electric motor24C.

FIG.5shows an example of output power distribution when both the first generator46aand the second generator46bare operating. In this case, the electric motor control section90controls each VTOL electric motor20V of the first drive system28and each cruise electric motor24C of the first drive system28to make the output power thereof equal to the output power command value. Further, the electric motor control section90controls each VTOL electric motor20V of the second drive system30and each cruise electric motor24C of the second drive system30to make the output power thereof equal to the output power command value.

FIG.6shows an example of output power distribution when the first generator46ais stopped and the second generator46bis operating. The electric motor control section90controls each VTOL electric motor20V of the first drive system28to make the output power thereof smaller than the output power command value. In addition, the electric motor control section90controls each VTOL electric motor20V of the second drive system30to make the output power thereof larger than the output power command value. As a result, the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30is increased by an amount corresponding to the decrease in the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28. As a result, the sum of the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28and the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30is maintained before and after the change of the output power distribution.

Further, the electric motor control section90controls each cruise electric motor24C of the first drive system28to set the output power thereof to 0. In addition, the electric motor control section90controls each cruise electric motor24C of the second drive system30to make the output power thereof larger than the output power command value. However, the sum of the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the first drive system28and the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the second drive system30is decreased as compared with that before the change of the output power distribution. Thus, the thrust generated by the cruise rotors22C decreases in the aircraft10as a whole. As a result, the airspeed of the aircraft10decreases, and the drag acting on the fuselage12can be reduced, whereby the power consumption rate can be improved. Therefore, the operation duration of the VTOL electric motors20V and the cruise electric motors24C of the first drive system28can be increased.

Advantageous Effects

For example, when the gas turbine44of the first main power source device32ais stopped, the first generator46acannot generate electric power. Even in this case, the VTOL electric motors20V and the cruise electric motors24C of the first drive system28can continue to operate by the batteries50of the first drive system28. It is conceivable to increase the capacity of the battery50in order to increase the flight duration of the aircraft10in a case where the generator46of one of the main power source devices32cannot generate electric power. However, in this case, the weight of the battery50increases and the weight of the fuselage12increases. Also, the cost of manufacturing the aircraft10increases.

In the flight controller80of the present embodiment, the electric motor control section90changes the output power distribution in the following case. The case indicates a state in which electric power cannot be supplied from the first generator46ato the VTOL electric motors20V and the cruise electric motors24C of the first drive system28and electric power can be supplied from the second generator46bto the VTOL electric motors20V and the cruise electric motors24C of the second drive system30. In this case, the VTOL electric motors20V and the cruise electric motors24C of the first drive system28are driven by the electric power of the batteries50of the first drive system28.

The output power distribution is changed in the following manner. Specifically, the electric motor control section90makes the output power of each VTOL electric motor20V of the first drive system28smaller than the output power command value, and makes the output power of each VTOL electric motor20V of the second drive system30larger than the output power command value. Further, the electric motor control section90makes the output power of each cruise electric motor24C of the first drive system28smaller than the output power command value, and makes the output power of the cruise electric motor24C of the second drive system30larger than the output power command value.

As a result, the power consumption of each VTOL electric motor20V and each cruise electric motor24C of the first drive system28decreases. Therefore, the operation duration of the VTOL electric motors20V and the cruise electric motors24C of the first drive system28by the electric power of the batteries50of the first drive system28can be increased, whereby the flight duration of the aircraft10can be increased.

In the flight controller80of the present embodiment, the electric motor control section90reduces the sum of the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the first drive system28and the thrust generated by the cruise rotors22C driven by the cruise electric motors24C of the second drive system30, as compared with that before the change of the output power distribution. Thus, the thrust generated by the cruise rotors22C decreases in the aircraft10as a whole. As a result, the airspeed of the aircraft10decreases, and the drag acting on the fuselage12can be reduced, whereby the power consumption rate can be improved. Therefore, the operation duration of the VTOL electric motors20V and the cruise electric motors24C of the first drive system28can be increased, and flight duration of the aircraft10can be increased.

In the aircraft10of the present embodiment, among the VTOL rotors18V disposed on the left side of the center line A of the fuselage12, the number of the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28is equal to the number of the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30. Further, among the VTOL rotors18V disposed on the right side of the center line A of the fuselage12, the number of the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28is equal to the number of the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30. As a result, when the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the first drive system28is decreased and the thrust generated by the VTOL rotors18V driven by the VTOL electric motors20V of the second drive system30is increased, the attitude of the fuselage12can be stabilized.

Note that the present invention is not limited to the above disclosure, and various modifications are possible without departing from the essence and gist of the present invention.

In the first embodiment, the first power supply circuit38aand the second power supply circuit38bare not connected to each other and are provided independently. However, the first power supply circuit38aand the second power supply circuit38bmay be connected to each other via a switch.

In the first embodiment, one cruise rotor22C is provided on the left side of the center line A of the fuselage12, and one cruise rotor22C is provided on the right side of the center line A of the fuselage12. On the other hand, two cruise rotors22C may be provided on the left side of the fuselage12, and two cruise rotors22C may be provided on the right side of the fuselage12. In this case, one cruise rotor22C is driven by one cruise electric motor24C. Further, one cruise rotor22C may be provided at the center of the fuselage12in the left-right direction. In this case, one cruise rotor22C is driven by two cruise electric motors24C.

The VTOL rotor18V and the cruise rotor22C of the first embodiment may be coaxial rotors. In this case, one of the two coaxial rotors may be driven by the electric motor of the first drive system28, and the other rotor may be driven by the electric motor of the second drive system30.

Invention Obtained from Embodiment

The invention that can be grasped from the above embodiment will be described below.

Provided is the control device (80) for the aircraft (10), the aircraft including: at least one first generator (46a) configured to generate electric power; at least one first battery (50a,50b) configured to store electric power; at least one first electric motor (20V1_1,20V4_1,20V5_1,20V8_1,24C1_1,24C3_1) configured to operate with electric power supplied from the first generator and the first battery; at least one second generator (46b) configured to generate electric power; at least one second battery (50c,50d) configured to store electric power; at least one second electric motor (20V2_2,20V3_2,20V6_2,20V7_2,24C2_2,24C4_2) configured to operate with electric power supplied from the second generator and the second battery; and the plurality of rotors (18V,22C) configured to generate thrust acting on the fuselage (12), wherein the control device includes the electric motor control section (90) configured to control the first electric motor and the second electric motor, and wherein each of the rotors is driven by one of the first electric motor or the second electric motor, or both of the first electric motor and the second electric motor, and in a case where supply of electric power from the first generator to the first electric motor is unavailable and supply of electric power from the second generator to the second electric motor is available, the electric motor control section reduces power consumption of the first electric motor by reducing thrust generated by the rotor driven by the first electric motor, and increases thrust generated by the rotor driven by the second electric motor, as compared with a case where the supply of the electric power from the first generator to the first electric motor is available and the supply of the electric power from the second generator to the second electric motor is available. According to this feature, the operation duration of the first electric motor can be increased, and the flight duration of the aircraft can be increased.

In the above-described control device for the aircraft, each of the first electric motor and the second electric motor may drive the horizontal rotor (22C) configured to generate thrust in the horizontal direction, and in the case where the supply of the electric power from the first generator to the first electric motor is unavailable and the supply of the electric power from the second generator to the second electric motor is available, the electric motor control section may reduce the sum of thrust generated by the horizontal rotor driven by the first electric motor and thrust generated by the horizontal rotor driven by the second electric motor, as compared with the case where the supply of the electric power from the first generator to the first electric motor is available and the supply of the electric power from the second generator to the second electric motor is available. According to this feature, the operation duration of the first electric motor can be increased, and the flight duration of the aircraft can be increased.

In the above-described control device for the aircraft, the number of the rotors disposed on one side of the center line of the fuselage in the left-right direction may be equal to the number of the rotors disposed on another side of the center, among the rotors disposed on the one side, the number of the rotors driven by the first electric motor may be equal to the number of the rotors driven by the second electric motor, and among the rotors disposed on the other side, the number of the rotors driven by the first electric motor may be equal to the number of the rotors driven by the second electric motor. According to this feature, the attitude of the fuselage can be stabilized.

In the above-described control device for the aircraft, each of the rotors may be the vertical rotor (18V) configured to generate thrust in the vertical direction or the horizontal rotor configured to generate thrust in the horizontal direction. According to this feature, the vertical rotor or the horizontal rotor can be driven for a longer period of time, and the flight duration of the aircraft can be increased.