Patent Description:
<CIT> 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.

In the technique disclosed in <CIT>, 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.

<FIG> is a schematic diagram of an aircraft <NUM>. The aircraft <NUM> of the present embodiment is an electric vertical take-off and landing aircraft (eVTOL aircraft). In the aircraft <NUM>, rotors are driven by electric motors. The aircraft <NUM> generates vertical thrust and horizontal thrust by the rotors. Further, the aircraft <NUM> is a hybrid aircraft. The aircraft <NUM> includes a generator and a battery as power sources of the electric motor. In the aircraft <NUM>, 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 aircraft <NUM> includes a fuselage <NUM>. The fuselage <NUM> is provided with a cockpit, a cabin, and the like. A pilot rides in the cockpit and controls the aircraft <NUM>. Passengers and the like ride in the cabin. The aircraft <NUM> may be automatically controlled.

The aircraft <NUM> includes a front wing <NUM> and a rear wing <NUM>. When the aircraft <NUM> moves forward, lift is generated in each of the front wing <NUM> and the rear wing <NUM>.

The aircraft <NUM> includes eight VTOL rotors 18V. The eight VTOL rotors 18V are a rotor 18V1, a rotor 18V2, a rotor 18V3, a rotor 18V4, a rotor 18V5, a rotor 18V6, a rotor 18V7, and a rotor 18V8.

The rotor 18V1, the rotor 18V3, the rotor 18V5, and the rotor 18V7 are disposed on the left side of a center line A of the fuselage <NUM> in the left-right direction. The rotor 18V2, the rotor 18V4, the rotor 18V6, and the rotor 18V8 are disposed on the right side of the center line A. That is, four VTOL rotors 18V are disposed on the left side of the center line A, and four VTOL rotors 18V are disposed on the right side of the center line A.

The center of gravity G of the aircraft <NUM> is located on the center line A of the fuselage <NUM>. In a state where the aircraft <NUM> is viewed from above, the center of gravity G is located between the rotor 18V4 and the rotor 18V6 in the front-rear direction of the fuselage <NUM>. Further, the center of gravity G is located between the rotor 18V3 and the rotor 18V5 in the front-rear direction of the fuselage <NUM>.

In a state where the aircraft <NUM> is viewed from above, the rotor 18V8 is provided at a position point-symmetrical to the rotor 18V1 with respect to the center of gravity G. The rotor 18V7 is provided at a position point-symmetrical to the rotor 18V2 with respect to the center of gravity G. The rotor 18V6 is provided at a position point-symmetrical to the rotor 18V3 with respect to the center of gravity G. The rotor 18V5 is provided at a position point-symmetrical to the rotor 18V4 with respect to the center of gravity G.

One VTOL electric motor 20V is provided for one VTOL rotor 18V. Specifically, an electric motor 20V1_1 is provided for the rotor 18V1, an electric motor 20V2_2 is provided for the rotor 18V2, an electric motor 20V3_2 is provided for the rotor 18V3, an electric motor 20V4_1 is provided for the rotor 18V4, an electric motor 20V5_1 is provided for the rotor 18V5, an electric motor 20V6_2 is provided for the rotor 18V6, an electric motor 20V7_2 is provided for the rotor 18V7, and an electric motor 20V8_1 is provided for the rotor 18V8. Each VTOL rotor 18V is driven by each VTOL electric motor 20V.

Each VTOL rotor 18V generates thrust mainly toward the upper side of the fuselage <NUM>. The thrust of each VTOL rotor 18V 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 rotor 18V increases as the output power of the VTOL electric motor 20V increases. Each VTOL rotor 18V 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 rotor 18V is used during attitude control.

By controlling the thrust of each VTOL rotor 18V, a propulsive force is applied mainly upward to the fuselage <NUM>. By controlling the thrust of each VTOL rotor 18V, a roll moment, a pitch moment, and a yaw moment are caused to act on the fuselage <NUM>. The VTOL rotor 18V corresponds to a vertical rotor of the present invention.

The aircraft <NUM> includes two cruise rotors 22C. The two cruise rotors 22C are a rotor 22C1 and a rotor 22C2. The rotor 22C1 and the rotor 22C2 are attached to the rear portion of the fuselage <NUM>. The rotor 22C1 is disposed on the left side of the center line A. The rotor 22C2 is disposed on the right side of the center line A. That is, one cruise rotor 22C is disposed on the left side of the center line A, and one cruise rotor 22C is disposed on the right side of the center line A.

Two cruise electric motors 24C are provided for each cruise rotor 22C. Specifically, an electric motor 24C1_1 and an electric motor 24C2_2 are provided for the rotor 22C1, and an electric motor 24C3_1 and an electric motor 24C4_2 are provided for the rotor 22C2. One cruise rotor 22C is driven by two cruise electric motors 24C.

Each cruise rotor 22C generates thrust mainly toward the front of the fuselage <NUM>. The thrust of each cruise rotor 22C 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 rotor 22C increases as the output power of the cruise electric motor 24C increases. Each cruise rotor 22C 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 rotor 22C, a propulsive force is applied mainly forward to the fuselage <NUM>. The cruise rotor 22C corresponds to a horizontal rotor of the present invention.

<FIG> is a schematic diagram showing a configuration of a power supply system <NUM>.

The aircraft <NUM> includes the electric motor 20V1_1, the electric motor 20V4_1, the electric motor 20V5_1, the electric motor 20V8_1, the electric motor 24C1_1, and the electric motor 24C3_1 as drive sources of a first drive system <NUM>. Hereinafter, the electric motor 20V1_1, the electric motor 20V4_1, the electric motor 20V5_1, and the electric motor 20V8_1 may be referred to as a VTOL electric motor 20V of the first drive system <NUM>. Further, the electric motor 24C1_1 and the electric motor 24C3_1 may be referred to as a cruise electric motor 24C of the first drive system <NUM>. The VTOL electric motor 20V of the first drive system <NUM> and the cruise electric motor 24C of the first drive system <NUM> correspond to a first electric motor of the present invention.

The aircraft <NUM> includes the electric motor 20V2_2, the electric motor 20V3_2, the electric motor 20V6_2, the electric motor 20V7_2, the electric motor 24C2_2, and the electric motor 24C4_2 as drive sources of a second drive system <NUM>. Hereinafter, the electric motor 20V2_2, the electric motor 20V3_2, the electric motor 20V6_2, and the electric motor 20V7_2 may be referred to as a VTOL electric motor 20V of the second drive system <NUM>. Further, the electric motor 24C2_2 and the electric motor 24C4_2 may be referred to as a cruise electric motor 24C of the second drive system <NUM>. The VTOL electric motor 20V of the second drive system <NUM> and the cruise electric motor 24C of the second drive system <NUM> correspond to a second electric motor of the present invention.

The power supply system <NUM> includes two main power source devices <NUM> and four auxiliary power source devices <NUM>. The power supply system <NUM> supplies electric power to four load modules <NUM>.

The two main power source devices <NUM> include a first main power source device 32a and a second main power source device 32b. The four auxiliary power source devices <NUM> include a first auxiliary power source device 34a, a second auxiliary power source device 34b, a third auxiliary power source device 34c, and a fourth auxiliary power source device 34d. The four load modules <NUM> include a first load module 36a, a second load module 36b, a third load module 36c, and a fourth load module 36d.

The power supply system <NUM> includes two power supply circuits <NUM>. The two power supply circuits <NUM> include a first power supply circuit 38a and a second power supply circuit 38b. The first power supply circuit 38a and the second power supply circuit 38b are not connected to each other and are provided independently.

Each power supply circuit <NUM> includes a main power source circuit <NUM> and an auxiliary power source circuit <NUM>. The main power source circuit <NUM> is provided for each main power source device <NUM>. The auxiliary power source circuit <NUM> is provided for each auxiliary power source device <NUM>.

Each main power source device <NUM> includes a gas turbine <NUM>, a generator <NUM>, and a power control unit (hereinafter referred to as PCU) <NUM>. The gas turbine <NUM> drives the generator <NUM>. As a result, the generator <NUM> generates electric power. The PCU <NUM> converts the AC power generated by the generator <NUM> into DC power and outputs the DC power to the main power source circuit <NUM>. When starting the gas turbine <NUM>, the PCU <NUM> converts the DC power supplied by the main power source circuit <NUM> into AC power, and outputs the AC power to the generator <NUM>. The generator <NUM> is operated by the AC power input from the PCU <NUM>, and the generator <NUM> drives the gas turbine <NUM>.

Hereinafter, the generator <NUM> in the first main power source device 32a may be referred to as a first generator 46a. Further, the generator <NUM> in the second main power source device 32b may be referred to as a second generator 46b.

Each auxiliary power source device <NUM> includes a battery <NUM>. The battery <NUM> is charged with DC power supplied from the main power source device <NUM>. Hereinafter, the battery <NUM> in the first auxiliary power source device 34a may be referred to as a first battery 50a. Further, the battery <NUM> in the second auxiliary power source device 34b may be referred to as a second battery 50b, the battery <NUM> in the third auxiliary power source device 34c may be referred to as a third battery 50c, and the battery <NUM> in the fourth auxiliary power source device 34d may be referred to as a fourth battery 50d.

The first battery 50a and the second battery 50b supply electric power to the VTOL electric motors 20V of the first drive system <NUM> and the cruise electric motors 24C of the first drive system <NUM>. That is, the first battery 50a and the second battery 50b function as power storage devices of the first drive system <NUM>. Hereinafter, each of the first battery 50a and the second battery 50b may be referred to as a battery <NUM> of the first drive system <NUM>. The third battery 50c and the fourth battery 50d supply electric power to the VTOL electric motors 20V of the second drive system <NUM> and the cruise electric motors 24C of the second drive system <NUM>. That is, the third battery 50c and the fourth battery 50d function as power storage devices of the second drive system <NUM>. Hereinafter, each of the third battery 50c and the fourth battery 50d may be referred to as a battery <NUM> of the second drive system <NUM>. The battery <NUM> of the first drive system <NUM> corresponds to a first battery of the present invention. The battery <NUM> of the second drive system <NUM> corresponds to a second battery of the present invention.

Each load module <NUM> includes two VTOL drive units <NUM> and one cruise drive unit <NUM>.

Each VTOL drive unit <NUM> includes an inverter <NUM> and the VTOL electric motor 20V. The inverter <NUM> converts the DC power supplied from the main power source circuit <NUM> into three phase AC power, and outputs the three phase AC power to the VTOL electric motor 20V.

The cruise drive unit <NUM> includes an inverter <NUM> and the cruise electric motor 24C. The inverter <NUM> converts the DC power supplied from the main power source circuit <NUM> into three phase AC power, and outputs the three phase AC power to the cruise electric motor 24C.

Each of the first load module 36a and the third load module 36c includes a converter <NUM>. The converter <NUM> steps down the voltage of the DC power supplied from the main power source device <NUM>, 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 PCU <NUM>, the inverters <NUM> and <NUM>, and the like.

Each main power source circuit <NUM> includes one common bus <NUM>, one contactor unit <NUM>, two contactor units <NUM>, one current sensor <NUM>, and two current sensors <NUM>.

The common bus <NUM> connects one main power source device <NUM> and two load modules <NUM>. The common bus <NUM> connects the two load modules <NUM> in parallel with the main power source device <NUM>.

The contactor unit <NUM> is provided between the main power source device <NUM> and the common bus <NUM>. The contactor unit <NUM> switches between a conduction state in which current flows between the main power source device <NUM> and the common bus <NUM>, and an interruption state in which the flow of the current between the main power source device <NUM> and the common bus <NUM> is interrupted. The contactor unit <NUM> includes a contactor 64a and a contactor 64b. The contactor 64a is provided on a positive line of the main power source circuit <NUM>. The contactor 64b is provided on a negative line of the main power source circuit <NUM>. The contactor unit <NUM> may include only one of the contactor 64a or the contactor 64b.

Each contactor unit <NUM> is provided between each load module <NUM> and the common bus <NUM>. The contactor unit <NUM> switches between a conduction state in which current flows between each load module <NUM> and the common bus <NUM>, and an interruption state in which the flow of the current between each load module <NUM> and the common bus <NUM> is interrupted. The contactor unit <NUM> includes a contactor 66a and a contactor 66b. The contactor 66a is provided on the positive line of the main power source circuit <NUM>. The contactor 66b is provided on the negative line of the main power source circuit <NUM>. The contactor unit <NUM> may include only one of the contactor 66a or the contactor 66b. When the contactor unit <NUM> includes only the contactor 64a, each contactor unit <NUM> preferably includes only the contactor 66b. When the contactor unit <NUM> includes only the contactor 64b, each contactor unit <NUM> preferably includes only the contactor 66a.

The current sensor <NUM> is provided between the contactor unit <NUM> and the common bus <NUM>. The current sensor <NUM> is provided on the positive line of the main power source circuit <NUM>. Each current sensor <NUM> is provided between each contactor unit <NUM> and the common bus <NUM>. Each current sensor <NUM> is provided on the positive line of the main power source circuit <NUM>.

Each auxiliary power source circuit <NUM> is connected to both the main power source circuit <NUM> and the load module <NUM>. The auxiliary power source circuit <NUM> supplies electric power from the auxiliary power source device <NUM> to the load module <NUM>. The auxiliary power source circuit <NUM> includes a contactor unit <NUM> and a current sensor <NUM>.

The contactor unit <NUM> is provided between the auxiliary power source device <NUM> and the load module <NUM>. The contactor unit <NUM> switches between a conduction state in which current flows between the auxiliary power source device <NUM> and the load module <NUM>, and an interruption state in which the flow of the current between the auxiliary power source device <NUM> and the load module <NUM> is interrupted. The contactor unit <NUM> includes a contactor 72a, a contactor 72b, and a precharge circuit 72c. The contactor 72a is provided on a positive line of the auxiliary power source circuit <NUM>. The contactor 72b is provided on a negative line of the auxiliary power source circuit <NUM>. The precharge circuit 72c is provided in parallel with the contactor 72b. The precharge circuit 72c includes a contactor 72d and a resistor 72e. The current sensor <NUM> is provided on the negative line of the auxiliary power source circuit <NUM>.

The contactor unit <NUM> may include only the contactor 72b and the precharge circuit 72c. The precharge circuit 72c may be provided in parallel with the contactor 72a. In this case, the contactor unit <NUM> may include only the contactor 72a and the precharge circuit 72c.

A diode <NUM> is provided between the main power source circuit <NUM> and each auxiliary power source circuit <NUM>. An anode of the diode <NUM> is connected to the main power source circuit <NUM>, and a cathode of the diode <NUM> is connected to the auxiliary power source circuit <NUM>. The diode <NUM> allows electric power to be supplied from the main power source circuit <NUM> to the auxiliary power source circuit <NUM>. The diode <NUM> prevents electric power from being supplied from the auxiliary power source circuit <NUM> to the main power source circuit <NUM>. When the main power source circuit <NUM> is short-circuited, electricity is prevented from flowing from the auxiliary power source device <NUM> to the main power source circuit <NUM>. As a result, even when the main power source circuit <NUM> is short-circuited, electric power can be supplied from the auxiliary power source device <NUM> to the load module <NUM>.

A transistor <NUM> is provided in parallel with the diode <NUM>. When the transistor <NUM> is ON, electric power is supplied from the auxiliary power source device <NUM> to the main power source circuit <NUM> while bypassing the diode <NUM>. The generator <NUM> is operated by the electric power supplied from the auxiliary power source device <NUM>, and the gas turbine <NUM> can be started.

Among the VTOL rotors 18V disposed on the left side of the center line A of the fuselage <NUM> (<FIG>), the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> are two rotors 18V1 and 18V5 (<FIG>). Among the VTOL rotors 18V disposed on the left side of the center line A of the fuselage <NUM> (<FIG>), the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> are two rotors 18V3 and 18V7 (<FIG>). That is, among the VTOL rotors 18V disposed on the left side of the center line A of the fuselage <NUM>, the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> is equal to the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM>.

Among the VTOL rotors 18V disposed on the right side of the center line A of the fuselage <NUM> (<FIG>), the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> are two rotors 18V4 and 18V8 (<FIG>). Among the VTOL rotors 18V disposed on the right side of the center line A of the fuselage <NUM> (<FIG>), the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> are two rotors 18V2 and 18V6 (<FIG>). That is, among the VTOL rotors 18V disposed on the right side of the center line A of the fuselage <NUM>, the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> is equal to the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM>.

The rotor 22C1 disposed on the left side of the center line A of the fuselage <NUM> is driven by the cruise electric motor 24C of the first drive system <NUM> and is driven by the cruise electric motor 24C of the second drive system <NUM> (<FIG>). That is, regarding the cruise rotors 22C disposed on the left side of the center line A of the fuselage <NUM>, the number of the cruise rotors 22C driven by the cruise electric motor 24C of the first drive system <NUM> is equal to the number of the cruise rotors 22C driven by the cruise electric motor 24C of the second drive system <NUM>.

The rotor 22C2 disposed on the right side of the center line A of the fuselage <NUM> is driven by the cruise electric motor 24C of the first drive system <NUM> and is driven by the cruise electric motor 24C of the second drive system <NUM> (<FIG>). That is, regarding the cruise rotors 22C disposed on the right side of the center line A of the fuselage <NUM>, the number of the cruise rotors 22C driven by the cruise electric motor 24C of the first drive system <NUM> is equal to the number of the cruise rotors 22C driven by the cruise electric motor 24C of the second drive system <NUM>.

The power supply system <NUM> includes a flight controller <NUM>. The flight controller <NUM> controls the thrust output from each VTOL rotor 18V and each cruise rotor 22C. <FIG> is a control block diagram of the flight controller <NUM>.

The flight controller <NUM> includes a computation section <NUM> and a storage section <NUM>. The computation section <NUM> is, for example, a processor such as a central processing unit (CPU) or a graphics processing unit (GPU). The computation section <NUM> includes an output power command value calculation section <NUM>, a main power source device monitoring section <NUM>, and an electric motor control section <NUM>. The output power command value calculation section <NUM>, the main power source device monitoring section <NUM>, and the electric motor control section <NUM> are realized by the computation section <NUM> executing programs stored in the storage section <NUM>. At least part of the output power command value calculation section <NUM>, the main power source device monitoring section <NUM>, and the electric motor control section <NUM> may 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 section <NUM>, the main power source device monitoring section <NUM>, and the electric motor control section <NUM> may be realized by an electronic circuit including a discrete device.

The storage section <NUM> is 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 section <NUM> may be included in the processor, the integrated circuit, or the like described above.

The output power command value calculation section <NUM> calculates an output power command value for each VTOL electric motor 20V, and an output power command value for each cruise electric motor 24C. 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 fuselage <NUM>, 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 aircraft <NUM> is 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 section <NUM> monitors the state of each main power source device <NUM>. For example, the main power source device monitoring section <NUM> monitors whether or not the generator <NUM> is operating, as the state of the main power source device <NUM>. When the generator <NUM> is stopped, electric power cannot be supplied from the generator <NUM> to each VTOL electric motor 20V and each cruise electric motor 24C.

The electric motor control section <NUM> controls each VTOL electric motor 20V to set the output power of each VTOL electric motor 20V to the power command value. The electric motor control section <NUM> controls each cruise electric motor 24C to set the output power of each cruise electric motor 24C to the power command value. When a predetermined condition is satisfied, the electric motor control section <NUM> changes the output power distribution in each VTOL electric motor 20V and each cruise electric motor 24C.

<FIG> is a flowchart showing a process of output power control. The output power control is performed by the electric motor control section <NUM>. The output power control is repeatedly executed at a predetermined cycle while the aircraft <NUM> is activated.

In step S1, the electric motor control section <NUM> determines whether or not both the first generator 46a and the second generator 46b are operating. When both the first generator 46a and the second generator 46b are operating, the process proceeds to step S2. When one of the first generator 46a or the second generator 46b is stopped, the process proceeds to step S3.

In step S2, the electric motor control section <NUM> controls each VTOL electric motor 20V and each cruise electric motor 24C 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 motor 20V and each cruise electric motor 24C becomes substantially equal to the power command value.

In step S3, the electric motor control section <NUM> determines whether the stopped generator is the first generator 46a or the second generator 46b. When the first generator 46a is stopped, the process proceeds to step S4. When the second generator 46b is stopped, the process proceeds to step S8.

When the first generator 46a is stopped, the electric motor control section <NUM> changes the output power distribution in steps S4 to S7.

In step S4, the electric motor control section <NUM> makes the output power of each VTOL electric motor 20V of the first drive system <NUM> smaller 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 motor 20V of the first drive system <NUM> decreases. In addition, if other conditions are constant, the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> decreases.

In step S5, the electric motor control section <NUM> makes the output power of each VTOL electric motor 20V of the second drive system <NUM> larger 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 motor 20V of the second drive system <NUM> increases. In addition, if the other conditions are constant, the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> increases.

In step S5, the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> is increased by an amount corresponding to the decrease in the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> in step S4. As a result, the sum of the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> and the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> is maintained before and after the change of the output power distribution.

In step S6, the electric motor control section <NUM> sets the output power of each cruise electric motor 24C of the first drive system <NUM> to <NUM>. Thereafter, the process proceeds to step S7. By the process of step S6, the power consumption of each cruise electric motor 24C of the first drive system <NUM> decreases. The thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the first drive system <NUM> becomes <NUM>.

In step S7, the electric motor control section <NUM> makes the output power of each cruise electric motor 24C of the second drive system <NUM> larger 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 motor 24C of the second drive system <NUM> increases. In addition, if the other conditions are constant, the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the second drive system <NUM> increases. The sum of the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the first drive system <NUM> and the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the second drive system <NUM> is decreased as compared with that before the change of the output power distribution.

When the second generator 46b is stopped, the electric motor control section <NUM> changes the output power distribution in steps S8 to S11.

In step S8, the electric motor control section <NUM> makes the output power of each VTOL electric motor 20V of the second drive system <NUM> smaller 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 motor 20V of the second drive system <NUM> decreases. In addition, if other conditions are constant, the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> decreases.

In step S9, the electric motor control section <NUM> makes the output power of each VTOL electric motor 20V of the first drive system <NUM> larger 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 motor 20V of the first drive system <NUM> increases. In addition, if the other conditions are constant, the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> increases.

In step S9, the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> is increased by an amount corresponding to the decrease in the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> in step S8. As a result, the sum of the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> and the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> is maintained before and after the change of the output power distribution.

In step S10, the electric motor control section <NUM> sets the output power of each cruise electric motor 24C of the second drive system <NUM> to <NUM>. Thereafter, the process proceeds to step S11. By the process of step S10, the power consumption of each cruise electric motor 24C of the second drive system <NUM> decreases. The thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the second drive system <NUM> becomes <NUM>.

In step S11, the electric motor control section <NUM> makes the output power of each cruise electric motor 24C of the first drive system <NUM> larger 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 motor 24C of the first drive system <NUM> increases. In addition, if the other conditions are constant, the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the first drive system <NUM> increases. The sum of the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the first drive system <NUM> and the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the second drive system <NUM> is decreased as compared with that before the change of the output power distribution.

<FIG> and <FIG> are image diagrams of output power distribution in each VTOL electric motor 20V and each cruise electric motor 24C. The numerical values of percentages shown in <FIG> and <FIG> indicate the ratio of the output power to the output power command value for each VTOL electric motor 20V and each cruise electric motor 24C.

<FIG> shows an example of output power distribution when both the first generator 46a and the second generator 46b are operating. In this case, the electric motor control section <NUM> controls each VTOL electric motor 20V of the first drive system <NUM> and each cruise electric motor 24C of the first drive system <NUM> to make the output power thereof equal to the output power command value. Further, the electric motor control section <NUM> controls each VTOL electric motor 20V of the second drive system <NUM> and each cruise electric motor 24C of the second drive system <NUM> to make the output power thereof equal to the output power command value.

<FIG> shows an example of output power distribution when the first generator 46a is stopped and the second generator 46b is operating. The electric motor control section <NUM> controls each VTOL electric motor 20V of the first drive system <NUM> to make the output power thereof smaller than the output power command value. In addition, the electric motor control section <NUM> controls each VTOL electric motor 20V of the second drive system <NUM> to make the output power thereof larger than the output power command value. As a result, the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> is increased by an amount corresponding to the decrease in the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM>. As a result, the sum of the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> and the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> is maintained before and after the change of the output power distribution.

Further, the electric motor control section <NUM> controls each cruise electric motor 24C of the first drive system <NUM> to set the output power thereof to <NUM>. In addition, the electric motor control section <NUM> controls each cruise electric motor 24C of the second drive system <NUM> to make the output power thereof larger than the output power command value. However, the sum of the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the first drive system <NUM> and the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the second drive system <NUM> is decreased as compared with that before the change of the output power distribution. Thus, the thrust generated by the cruise rotors 22C decreases in the aircraft <NUM> as a whole. As a result, the airspeed of the aircraft <NUM> decreases, and the drag acting on the fuselage <NUM> can be reduced, whereby the power consumption rate can be improved. Therefore, the operation duration of the VTOL electric motors 20V and the cruise electric motors 24C of the first drive system <NUM> can be increased.

For example, when the gas turbine <NUM> of the first main power source device 32a is stopped, the first generator 46a cannot generate electric power. Even in this case, the VTOL electric motors 20V and the cruise electric motors 24C of the first drive system <NUM> can continue to operate by the batteries <NUM> of the first drive system <NUM>. It is conceivable to increase the capacity of the battery <NUM> in order to increase the flight duration of the aircraft <NUM> in a case where the generator <NUM> of one of the main power source devices <NUM> cannot generate electric power. However, in this case, the weight of the battery <NUM> increases and the weight of the fuselage <NUM> increases. Also, the cost of manufacturing the aircraft <NUM> increases.

In the flight controller <NUM> of the present embodiment, the electric motor control section <NUM> changes the output power distribution in the following case. The case indicates a state in which electric power cannot be supplied from the first generator 46a to the VTOL electric motors 20V and the cruise electric motors 24C of the first drive system <NUM> and electric power can be supplied from the second generator 46b to the VTOL electric motors 20V and the cruise electric motors 24C of the second drive system <NUM>. In this case, the VTOL electric motors 20V and the cruise electric motors 24C of the first drive system <NUM> are driven by the electric power of the batteries <NUM> of the first drive system <NUM>.

The output power distribution is changed in the following manner. Specifically, the electric motor control section <NUM> makes the output power of each VTOL electric motor 20V of the first drive system <NUM> smaller than the output power command value, and makes the output power of each VTOL electric motor 20V of the second drive system <NUM> larger than the output power command value. Further, the electric motor control section <NUM> makes the output power of each cruise electric motor 24C of the first drive system <NUM> smaller than the output power command value, and makes the output power of the cruise electric motor 24C of the second drive system <NUM> larger than the output power command value.

As a result, the power consumption of each VTOL electric motor 20V and each cruise electric motor 24C of the first drive system <NUM> decreases. Therefore, the operation duration of the VTOL electric motors 20V and the cruise electric motors 24C of the first drive system <NUM> by the electric power of the batteries <NUM> of the first drive system <NUM> can be increased, whereby the flight duration of the aircraft <NUM> can be increased.

In the flight controller <NUM> of the present embodiment, the electric motor control section <NUM> reduces the sum of the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the first drive system <NUM> and the thrust generated by the cruise rotors 22C driven by the cruise electric motors 24C of the second drive system <NUM>, as compared with that before the change of the output power distribution. Thus, the thrust generated by the cruise rotors 22C decreases in the aircraft <NUM> as a whole. As a result, the airspeed of the aircraft <NUM> decreases, and the drag acting on the fuselage <NUM> can be reduced, whereby the power consumption rate can be improved. Therefore, the operation duration of the VTOL electric motors 20V and the cruise electric motors 24C of the first drive system <NUM> can be increased, and flight duration of the aircraft <NUM> can be increased.

In the aircraft <NUM> of the present embodiment, among the VTOL rotors 18V disposed on the left side of the center line A of the fuselage <NUM>, the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> is equal to the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM>. Further, among the VTOL rotors 18V disposed on the right side of the center line A of the fuselage <NUM>, the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> is equal to the number of the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM>. As a result, when the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the first drive system <NUM> is decreased and the thrust generated by the VTOL rotors 18V driven by the VTOL electric motors 20V of the second drive system <NUM> is increased, the attitude of the fuselage <NUM> can be stabilized.

Note that the present invention is not limited to the above disclosure, and various modifications are possible without departing from the scope defined by the appended claims.

In the first embodiment, the first power supply circuit 38a and the second power supply circuit 38b are not connected to each other and are provided independently. However, the first power supply circuit 38a and the second power supply circuit 38b may be connected to each other via a switch.

In the first embodiment, one cruise rotor 22C is provided on the left side of the center line A of the fuselage <NUM>, and one cruise rotor 22C is provided on the right side of the center line A of the fuselage <NUM>. On the other hand, two cruise rotors 22C may be provided on the left side of the fuselage <NUM>, and two cruise rotors 22C may be provided on the right side of the fuselage <NUM>. In this case, one cruise rotor 22C is driven by one cruise electric motor 24C. Further, one cruise rotor 22C may be provided at the center of the fuselage <NUM> in the left-right direction. In this case, one cruise rotor 22C is driven by two cruise electric motors 24C.

The VTOL rotor 18V and the cruise rotor 22C 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 system <NUM>, and the other rotor may be driven by the electric motor of the second drive system <NUM>.

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

Provided is the control device (<NUM>) for the aircraft (<NUM>), 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 (<NUM>), wherein the control device includes the electric motor control section (<NUM>) 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.

Claim 1:
A control device (<NUM>) for an aircraft (<NUM>),
the aircraft comprising:
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
a plurality of rotors (18V, 22C) configured to generate thrust acting on a fuselage (<NUM>),
wherein the control device comprises an electric motor control section (<NUM>) 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, characterised in that the electric motor control section (<NUM>) is configured to,
in a case that the electric motor control section (<NUM>) determines that 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, reduce power consumption of the first electric motor by reducing thrust generated by the rotor driven by the first electric motor, and increase 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.