An air mobility apparatus having redundancy of multiple power sources

Abstract

An air mobility implementing redundancy of multiple power sources includes: an actuator including a first propulsion assembly; a power source unit including a first battery having self-generation capability and a second battery that is rechargeable and outputs power determined based on a flight mode; and a processor controlling the actuator and the power source unit. The processor is configured to: in response to a failure state of the first battery, supply power of the second battery module to the first propulsion assembly, the supply power being determined based on the flight mode, and in response to a failure state of the second battery, supply power from the first battery to the first propulsion assembly, where the power is greater than or equal to power in a normal state.

Description

AN AIR MOBILITY APPARATUS HAVING REDUNDANCY OF MULTIPLE POWER SOURCES

[0001] The present disclosure relates to a mobility apparatus that implements redundancy of multiple power sources, and more specifically, to an air mobility apparatus that ensures flight safety by efficiently responding to failures of part of the multiple power sources.

[0002] Aircraft may provide rapid transportation of people and goods in densely populated urban areas or short-distance routes. Aircraft operating in urban environments may operate with low noise and eco-friendliness. For example, aircraft using eco-friendly energy may adopt powertrain systems powered by electric batteries.

[0003] In some cases, an aircraft powertrain system may be designed to include the redundancy and specific power of the power source. These considerations arise because takeoff and landing flights consume more power than cruising flights, and to increase flight range for practicality. In some cases, battery-based powertrains, which inherently have lower specific power compared to internal combustion engines, are limited in terms of power management during takeoff and landing, as well as in increasing flight range. In some cases, battery-based power sources are prone to failures caused by various factors.

[0004] The present disclosure describes an air mobility apparatus that implements redundancy in multiple power sources to ensure flight safety by efficiently responding to failures of part of the power sources.

[0005] According to one aspect of the subject matter described in this application, an air mobility implements redundancy of multiple power sources. The air mobility apparatus includes an actuator comprising a first propulsion assembly, a power source unit comprising (i) a first battery having self-generation capability and (ii) a second battery that is rechargeable and configured to output power determined based on a flight mode, the flight mode comprising a high-power mode and a low-power mode, and at least one processor configured to control the actuator and the power source unit. The least one processor is configured to, in response to a failure state of the first battery, supply power from the second battery to the first propulsion assembly, the power being determined based on the flight mode, and in response to a failure state of the second battery, supply power from the first battery to the first propulsion assembly, where the power supplied from the first battery is greater than or equal to a normal power output from the second battery in a normal state of the second battery.

[0006] It is possible to provide an air mobility apparatus that implements redundancy to ensure flight safety by efficiently responding to failures of part of the multiple power sources.

[0007] FIG. 1 is a diagram schematically illustrating an example of modules of a mobility apparatus.

[0008] FIG. 2 is a diagram illustrating an example of an air mobility apparatus.

[0009] FIG. 3 is a diagram illustrating an example of an air mobility apparatus implementing a hovering mode.

[0010] FIG. 4 is a diagram illustrating an example of an upper portion of an air mobility apparatus that implements a hovering mode.

[0011] FIG. 5 is a diagram illustrating an example of an air mobility apparatus implementing a cruising mode.

[0012] FIG. 6 is a diagram schematically illustrating an example of modules of an air mobility apparatus.

[0013] FIG. 7 is a diagram illustrating an example of an air mobility apparatus.

[0014] FIG. 8 is a diagram illustrating an example of a layout of an energy supply architecture between a power source unit and a propulsion assembly.

[0015] FIG. 9 is a diagram illustrating an example of a layout of an energy supply architecture between a power source unit and a propulsion assembly.

[0016] FIG. 10 is a diagram illustrating an example connection between the power source unit and the propulsion assembly.

[0017] FIG. 11 is a diagram illustrating an example of an electric circuit of the power source unit.

[0018] FIG. 12 is a diagram illustrating an example of an electric circuit of the propulsion assembly.

[0019] FIG. 13 is a diagram illustrating an example profile of the first propulsion assembly by flight mode.

[0020] FIG. 14 is a diagram illustrating an example profile of the second propulsion assembly by flight mode.

[0021] FIG. 15 is a diagram showing an example of power supply to the propulsion assembly in high-power mode with three multiple power sources operating in a normal state.

[0022] FIG. 16 is a diagram showing an example of power supply to the propulsion assembly in low-power mode with three multiple power sources operating in a normal state.

[0023] FIG. 17 is a diagram showing an example of redundancy control responding to failure in three multiple power sources.

[0024] FIG. 18 is a diagram showing an example operation of a bidirectional converter.

[0025] FIG. 19 is a diagram showing an example of redundancy control responding to failure in three multiple power sources.

[0026] FIG. 20 is a diagram showing an example of redundancy control responding to failure in three multiple power sources.

[0027] FIG. 21 is a diagram showing an example of power supply to the propulsion assembly in high-power mode with two multiple power sources operating in a normal state.

[0028] FIG. 22 is a diagram showing an example of power supply to the propulsion assembly in low-power mode with two multiple power sources operating in a normal state.

[0029] FIG. 23 is a diagram showing an example of redundancy control responding to failure in two multiple power sources.

[0030] FIG. 24 is a diagram showing an example of redundancy control responding to failure in two multiple power sources.

[0031] Hereinafter, one or more implementations of the present disclosure will be described in detail with reference to the accompanying drawings so that those skilled in the art may easily implement the present disclosure. However, the present disclosure may be implemented in various different ways, and is not limited to the implementations described therein.

[0032] Hereinafter, FIG. 1 is a schematic diagram showing an example of modules of a mobility apparatus. The modules may be used in various types of mobility apparatuses.

[0033] In some implementations, the mobility apparatus 10 may be a moving body with mobility in the present disclosure. Mobility may involve moving from one point to a specific point while loading people, objects, and / or cargo for specific purposes. The mobility apparatus 10 may move for transportation and other purposes. Other purposes may include, for example, detecting or monitoring the environment around the mobility apparatus 10 by mounting an observation device. In some examples, the mobility apparatus 10 may be equipped with a camera to capture or analyze the surrounding environment and transmit the captured or analyzed image to a designated device. The mobility apparatus 10 may be used for various purposes beyond the aforementioned examples.

[0034] The mobility apparatus 10 may move through different spaces such as land, underground, air, space, sea, and / or underwater. Land or underground mobility apparatuses 10 may be provided, for example, in the form of vehicles or robots. Air or space mobility apparatuses 10 may be provided, for example, in the form of air mobility apparatuses, such as fixed-wing or rotary-wing aircraft, AAM (Advanced Air Mobility) actively developed recently, unmanned aerial vehicles, drones, rockets, or vehicles mounted on satellites. Sea or underwater mobility apparatuses 10 may include, for example, ships or submarines. The mobility apparatus 10 may move across multiple spaces without being limited to a specific space, such as amphibious vehicles or flying vehicles.

[0035] Furthermore, the mobility apparatus 10 may be moved manually, autonomously, by remote control or by a combination of these methods. Manual operation may be implemented by an operator or pilot using an interface, such as a control device provided in the mobility apparatus 10 or by remote control from a control center or an external control station. Autonomous control, that is, autonomous movement, may be performed by the independent processing of the mobility apparatus 10 or a combination of remote control through the control center and the cooperation between the mobility apparatus 10 and the control center. Combinations of the above-mentioned controls may be implemented, for example, depending on the movement control plan and the movement situation. In a movement control plan for an air mobility apparatus, takeoff and landing may be performed by remote control or manual operation, and cruising flights may be operated by autonomous control. The flight under normal travel conditions in the air mobility apparatus may be performed by autonomous control, while the flight in emergency situations may be controlled by remote control or manual operation. An emergency situation may be, for example, an abnormal condition of the power source unit 18 or the actuating unit 20 of the mobility apparatus, a sudden change in flight behavior due to adverse weather conditions or the sudden appearance of an obstacle, etc.

[0036] In some examples, the mobility apparatus 10, which operates in various forms, may be designed differently depending on the use, movement space, driving method, control method, and other factors. However, from a comprehensive perspective of mobility, it may have common functional modules, as illustrated in FIG. 1. FIG. 1 describes the common functions in various types of mobility apparatuses 10. Accordingly, unique functional modules utilized in each type are omitted, but the implementations of the present disclosure do not exclude the modules omitted in FIG. 1, nor are they excluded from the scope of the present disclosure.

[0037] The mobility apparatus 10 may include a sensor unit 12, a communication unit 14, and a load device 16.

[0038] The sensor unit 12 may be equipped with various types of detectors that detect the various states and situations occurring in the external and internal environments of the mobility apparatus 10. The sensor unit 12 includes a positioning sensor to identify the location information of the mobility apparatus 10. That is, the sensor unit 12 may include various heterogeneous sensors and acquire sensing data detected from each sensor. The sensor unit 12 may acquire sensor data used for movement control, status data detecting the state of modules constituting the mobility apparatus 10, situation data detecting the situation of passengers and / or cargo. The sensor unit 12 may provide the data to a processor 26 that triggers specific functions and actions. In the present disclosure, movement control may include at least one of the following: linear movement, turning, acceleration, deceleration, attitude control of the mobility apparatus 10, braking, and hovering. In the present disclosure, hovering may be control that generates thrust toward the downward or vertical direction relative to the mobility apparatus 10 to cause specific movements or movements of the mobility apparatus 10. The specific action or movement may include, for example, takeoff, landing, or substantial stationary flight within a limited range. The data from the sensor unit 12 described above is merely exemplary and may further include sensor data that detects various situations not enumerated here.

[0039] The communication unit 14 may support mutual communication with other devices to exchange data with external devices. Other devices may include, for example, a server controlling the mobility apparatus 10 or exchanging data related to the movement control of the mobility apparatus 10, ancillary devices supporting movement, and other mobility apparatuses. The server may be referred to by various terms, such as control device, management device, control station, and cloud server. The communication unit 14 may transmit data generated or stored during movement to other devices and receive data and software modules transmitted from other devices. The protocol applied to the communication unit 14 may be determined according to the type of mobility apparatus 10, and the communication unit 14 may communicate with other vehicles or other devices based on cellular communication, WAVE (Wireless Access in Vehicular Environment) communication, DSRC (Dedicated Short Range Communication), near-field communication or on other communication methods. In some examples, the communication unit 14 may include a transmitter, a receiver, or a transceiver configured to communicate signals wirelessly or via wires. The aforementioned communication protocols and methods are merely examples and are not limited to these.

[0040] The load device 16 is installed on the mobility apparatus 10 and may be an auxiliary device that consumes power supplied from the power source unit 18 or converted from the output of the power source unit 18 by a command for use by the user or management of the load. The load device 16 in the present disclosure may be a type of non-mobility electrical device, excluding the mobility power system used in the driving unit 22. The load device 114 may include, for example, a display system, an air conditioning system, a lighting system, a seat system, and various devices installed in the mobility apparatus 10.

[0041] In some implementations, the mobility apparatus 10 may include an interface that receives requests for operations of the movement control and the load device 16. The interface may be implemented as a hardware device or a software interface. The hardware interface may be a hardware control device for movement operations by the user for the mobility apparatus 10, such as a control stick for aviation, a steering wheel for land vehicles, pedals for land vehicles, buttons, a rudder for marine vessels, etc., but is not limited to these. The software interface may include, for example, a graphical user interface (GUI) of a touch-sensitive display, but is not limited to this.

[0042] The mobility apparatus 10 may also include a power source unit 18, an actuating unit 20, and a driving unit 22.

[0043] The power source unit 18 may generate and supply power and electricity used for the mobility power system, such as the driving unit 22, and for the load device 16. The mobility apparatus 10 may generate energy using at least one of various energy sources. When the mobility apparatus 10 is driven based on electric energy, the power source unit 18 may include, for example, an electric battery or a combination of an electric battery and a charging module that charges the battery. In some cases, where the power source unit 18 consists solely of an electric battery, the electric battery may be charged at a charging station or by another mobility apparatus to supply power. When the power source unit 18 is a combination of an electric battery and a charging module, the charging module may employ at least one of a fuel cell and an engine based on fossil energy. The fuel cell may use substances such as hydrogen gas to produce electricity. In the case of an engine, the power source unit 18 may include a generator coupled with the engine, and the generator may convert mechanical energy generated by the engine into electrical energy to charge the electric battery.

[0044] In some examples, where the mobility apparatus 10 is driven based on fossil energy or nuclear fuel, the power source unit 18 may be configured with an internal combustion engine, turbine engine, or nuclear fuel-based engine. The mobility apparatus 10 may also have a hybrid-type power source unit 18 including an engine based on fossil energy and an electric battery. The hybrid-type power source unit 18 may charge the electric battery using the output of the engine generated during movement, and depending on the state of the mobility apparatus 10, the characteristics of the movement path, or the movement situation, it may select either the power from the engine or the electric battery to generate the driving force of the driving unit 22. In another example, the hybrid-type power source unit 18 may include an electric battery that can be charged by an external source and an engine. The processor 26 may switch between the engine power and the electric battery power depending on various situations and states to apply energy to the driving unit 22, thereby generating driving force.

[0045] The actuating unit 20 may include at least one module that implements movement operations. In some examples, where the mobility apparatus 10 is an air mobility apparatus, the actuating unit 20 may include mechanical and software components that perform at least one of the following operations: flight attitude control, hovering control related to takeoff and landing, altitude change control, and turning operation control. The flight attitude may relate to the roll, yaw, and pitch of the air mobility apparatus. If the mobility apparatus 10 is a ground mobility apparatus, the actuating unit 20 may include mechanical and software components that realize at least one of the following driving operations: longitudinal control such as acceleration and deceleration, lateral control such as steering. In the present disclosure, the actuating unit 20 may also be referred to as an actuator.

[0046] The actuating unit 20 may include the driving unit 22. The driving unit 22 is a module that implements external operations such as linear movement, turning, acceleration, deceleration, attitude control of the mobility apparatus 10, braking, and hovering. The driving unit 22 may be implemented in various forms depending on the type of mobility apparatus 10.

[0047] In some examples, where the mobility apparatus 10 is an air mobility apparatus, the driving unit of the fixed-wing air mobility apparatus may be a turbine engine, flap installed on the main wing or tail wing and related to operations such as thrust and lift. In another example, the fixed-wing air mobility apparatus may additionally include a propulsion assembly such as a propeller installed on a specific part of the main wing. The driving unit of a rotary-wing air mobility apparatus may include a rotor-type propulsion assembly and flaps. installed on the upper part of the fuselage and the tail wing. Depending on the specifications, air mobility apparatuses may also include landing gear such as wheels for takeoff and landing, which may be housed within the fuselage during flight. The driving unit 22 of an AAM-type air mobility apparatus may be equipped with a rotor-type propulsion assembly similarly to a rotary-wing air mobility. The propulsion assembly applied to the AAM-type air mobility apparatus may be fixed at least non-tiltable to the main wing, or may be installed at least tiltably to the main wing. In another example, the propulsion assembly applied to the AAM-type air mobility apparatus may be installed multiple times inside the main wing. Additionally, the driving unit of the AAM-type air mobility apparatus may be configured to rotate the wing to which the propulsion assembly is coupled within a certain angle range. Depending on the specifications, the driving unit of the AAM-type air mobility apparatus may be provided with wheels, such as landing gear, that are accommodated within the fuselage during flight and extended during takeoff and landing. If the AAM-type air mobility apparatus is driven based on electric energy, the driving unit 22 may include a motor and inverter that primarily rotate the propeller with electricity. If the AAM-type air mobility apparatus is driven based on non-electric energy, such as fossil energy, the driving unit 22 may primarily include modules that transmit the rotational power generated by the internal combustion engine to the propeller.

[0048] In some examples, where the mobility apparatus 10 is a ground mobility apparatus, the driving unit 22 may include a plurality of wheels, a driving force transmission module for generating driving force and applying or transmitting driving force to the wheels, a braking module for decelerating the driving of the wheels, and a steering module for realizing lateral control of the wheels. The wheel, the driving force transmission modules, the braking modules, etc., may form a driving assembly, and driving assemblies may be provided depending on the number of wheels. If the ground mobility apparatus is driven based on electric energy, the driving force transmission module may including a motor module that generates driving force based on the electric power output from the electric battery. If the ground mobility apparatus is driven based on fossil energy, the driving force transmission module may include transmission and gear module that transmits power from the internal combustion engine.

[0049] In some examples, where the mobility apparatus 10 is a sea or underwater mobility apparatus, the driving unit 22 may include a rudder, propulsion propeller, and modules that transmit power and specific motions to these components.

[0050] The mobility apparatus 10 may also include memory 24 and a processor 26.

[0051] The memory 24 stores applications and various data for controlling the mobility apparatus 10 and may load applications or read / write data in response to a request from the processor 26. Applications and data vary depending on the type and detailed specifications of the mobility apparatus 10 and may include sensor data related to movement control, state data related to the movement control of the mobility apparatus, data received from other devices, and data related to energy control between the power source unit 18 and the driving unit 22. Additionally, applications and data may include data related to modules responsible for functions other than control, software related to the operation of the mobility computing system, information and applications for autonomous movement, path information, and various information and control programs for user convenience.

[0052] In relation to the present disclosure, the processor 26 may use the applications, instructions, and data stored in the memory 24 to handle movement control, path control, energy control, control of the load device 16, autonomous movement control, and convenience functions. The processor 26 may also have different control processes depending on the type and detailed specifications of the mobility apparatus 10. The processor 26 may be implemented as a single processing module. Alternatively, in some implementations, the processing according to the above-mentioned matters may be distributed across multiple processing modules, and the processor 26 may collectively refer to multiple processing modules in the present disclosure. That is, the process described in this application may be performed by at least one processor. In some examples, the at least one processor may include an electric circuit, application specific integrated circuits (ASICs), Digital signal processors (DSPs), digital signal processing devices (DSPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), general processors, controllers, microcontrollers, microprocessors, artificial intelligence chips, etc.

[0053] Hereinafter, the air mobility apparatus related to the implementation of the present disclosure will be described. Specifically, the present disclosure will primarily describe the AAM-type air mobility apparatus among various types of air mobility apparatuses. Although the AAM-type air mobility apparatus is described, the function, the modules and the apparatus described in the present disclosure may also be applied to other types of air mobility apparatuses if they are technically compatible. Broadly speaking, if the function, the modules and the apparatus described in the present disclosure are technically compatible, they may also be applied to ground, underground, space, sea, and underwater mobility apparatuses.

[0054] The AAM-type air mobility apparatus may be classified into types such as VTOL (Vertical Take-off and Landing), STOL (Short Take-off and Landing), and CTOL (Conventional Take-off and Landing) according to the take-off and landing method. VTOL refers to an air mobility apparatus that can take off and land vertically without a runway, STOL refers to an air mobility apparatus that can take off and land on a relatively short runway or at low speed, and CTOL refers to an air mobility apparatus that requires a longer runway or higher speed than STOL for take-off and landing. Additionally, VTOL may be further classified into vectored thrust (or tilting rotor) type, lift-and-cruise type, and multicopter type according to the propulsion method. In the tilting rotor type, at least portion of plurality of propulsion assemblies may be tilted to be fully utilized for both hovering and cruising. In the present disclosure, flight control according to the hovering mode may generate thrust in the downward or vertical direction relative to the fuselage of the air mobility apparatus, thereby causing a specific movement in the mobility apparatus 10. The specific movement may include, for example, takeoff, landing, or substantially stationary flight within a limited range. Flight control according to the cruising mode may relate to control related to forward flight or horizontal flight. In the lift-and-cruise type, the propulsion assemblies for hovering and cruising may be provided separately without tilting the rotors according to the flight mode. The multicopter type has multiple propulsion assemblies arranged horizontally relative to the fuselage similarly to a drone in that, and it can realize takeoff, landing, and cruising flight through attitude control of the fuselage.

[0055] FIG. 2 is a diagram illustrating an air mobility apparatus according to one example of the present disclosure.

[0056] The air mobility apparatus 30 of one example may include a streamlined fuselage 32, a main wing 34 arranged on the left and right sides of the central region of the fuselage 32, and a tail wing 36 arranged at the end region of the fuselage 32. The fuselage 32 may include, for example, a cabin for the pilot and passengers and space for loading cargo. The main wing 34 and the tail wing 36 may be configured as fixed wings.

[0057] The main wing 34 may have an inboard region and an outboard region. The main wing tilting propulsion assembly 38 with a tilting rotor may be arranged in the front region of the left and right outboard areas. The main wing non-tilting propulsion assemblies 40a and 40b , arranged horizontally relative to the fuselage 32, may be located in the inboard region and the rear region of the outboard areas, and may be referred to as lift propulsion assemblies. Additionally, the tail wing tilting propulsion assembly 42 with a tilting rotor may be arranged at the end of the tail wing.

[0058] Each of the propulsion assemblies described above may include a propeller with multiple blades, a hub to which the ends of each blade are connected, a motor that supplies rotational force to the shaft, an inverter that adjusts the power of the power source unit 18 based on flight conditions and motor specifications, and a propulsion assembly cooling module that cools the motor and inverter.

[0059] The main wing tilting propulsion assembly 38 and the tail wing tilting propulsion assembly 42 may include a nacelle that accommodates and covers the aforementioned components. The main wing tilting propulsion assembly 38 and the tail wing tilting propulsion assembly 42 may be fixed and may be tilted with respect to the boom by hinge coupling between nacelle and boom protruded from the outboard area of the main wing and the end of the tail wing. In the present disclosure, the propulsion assembly may also be referred to as a rotor propulsion assembly. When the flight mode is a hovering mode, such as during takeoff and landing, the main wing and tail wing tilting propulsion assemblies 38 and 42 may be tilted by the hinge coupling so that the blade arrangement direction is substantially parallel to that of the main wing non-tilting propulsion assembly 40a. When the flight mode is a cruising mode, such as during horizontal flight, the main wing and tail wing tilting propulsion assemblies 38 and 42 may be tilted by the hinge coupling so that the blade arrangement direction is at a certain angle relative to that of the main wing non-tilting propulsion assemblies 40a and 40b. FIG. 2 illustrates an assembly for tilting, utilizing hinge coupling between the mount and nacelle. In another example, the tilting assembly may be implemented using a different mechanism that does not rely on hinge coupling. For example, a multi-joint linkage may connect respective predetermined point of the nacelle, which houses the modules of the main wing non-tilting propulsion assemblies 40a and 40b, and the mount, and then the nacelle and the mount may not have a physical coupling point between them. This would allow the main wing and tail wing tilting propulsion assemblies 38 and 42 to pivot substantially perpendicularly to the fuselage 32 during hovering mode as the multi-joint linkage extends. In cruising mode, the multi-joint linkage contracts, allowing the main wing and tail wing tilting propulsion assemblies 38 and 42 to pivot substantially parallel to the fuselage 32. The tilting of the main wing and tail wing tilting propulsion assemblies 38 and 42 may be synchronized according to the flight mode, and the tilting of each propulsion assembly may be adjusted differently according to attitude control and flight conditions within the same flight mode.

[0060] The main wing non-tilting propulsion assemblies 40a and 40b may be arranged at the ends of the boom extending from the front area of the inboard section and the rear area of the outboard section of the main wing 34. The main wing non-tilting propulsion assemblies 40a and 40b may be formed as dual propellers that overlap each other, or as a single propeller, as illustrated in FIG. 2. The main wing non-tilting propulsion assemblies 40a and 40b are primarily activated during hovering mode, but they may also be activated according to attitude control and flight conditions even during cruising mode. In the present disclosure, the non-tilting propulsion assemblies 40a and 40b may also be referred to as lift rotors or lift assemblies.

[0061] Although the present disclosure describes the tilting and non-tilting propulsion assemblies as being arranged as shown in FIG. 2, the arrangement of these propulsion assemblies is not limited to the example shown in FIG. 2 and the propulsion assemblies may be arranged in different locations. Furthermore, although FIG. 2 depicts four tilting propulsion assemblies and two non-tilting propulsion assemblies, the number of units is not limited to these, and the number may vary. Variations in the arrangement and number of tilting and non-tilting propulsion assemblies are illustrated in FIGS. 3 to 5, but other forms not mentioned in the present disclosure may also be designed.

[0062] With reference to FIGS. 3 to 5, the air mobility apparatus will be described. FIG. 3 is a diagram illustrating an air mobility apparatus implementing a hovering mode. FIG. 4 is a diagram showing the upper part of the air mobility apparatus implementing the hovering mode. FIG. 5 is a diagram illustrating an air mobility apparatus implementing a cruising mode.

[0063] In some implementations, the air mobility apparatus, as shown in FIGS. 3 to 5, is substantially similar to the example in FIG. 2 except for the detailed configuration of the tilting propulsion assemblies. Therefore, the description of the same functions and components as in FIG. 2 is omitted, and the description will focus on the differences in the air mobility apparatus.

[0064] In another example, the air mobility apparatus 50 may include a fuselage 52, a main wing 54 arranged on the left and right sides of the central region of the fuselage 52, and a tail wing 56 arranged at the end region of the fuselage 52.

[0065] The main wing 54 may have an inboard region and an outboard region. The tilting propulsion assemblies 58 with tilting rotors may be arranged in the front region of the left and right outboard areas. The tilting propulsion assemblies 58 may be fixed by the boom prepared at the outboard area of the main wing. The non-tilting propulsion assemblies 60 and 64, arranged horizontally relative to the fuselage 52, may be located in the front and rear areas of the rod extended from the right and left inboard region, and they may be referred to as lift propulsion assemblies.

[0066] In some examples, the outboard area of the main wing 54 may be partitioned from the horizontal surface 54a of the inboard area and have the lift surface 54b, and the tilting propulsion assemblies 58 may be mounted on the boom protruding from the lift surface. The lift surface 54b may have a winglet bent at the end of the wing. In one example, the tilting propulsion assemblies 58 and the lift surface 54b may be configured to rotate or move together. If the lift surface 54b may be structurally separated from the boom like another example different from Figs 3 to 5, the tilting propulsion assemblies 58 may move or rotate together with the lift surface except for a part of the lift surface including the fuselage 52. In yet another example, only the hub and blades of the tilting propulsion assemblies 58 may be operated to move or rotate.

[0067] According to the details described above, when the flight mode is a hovering mode, the tilting propulsion assemblies 58 may be tilted to be substantially parallel to the blade arrangement direction of the non-tilting propulsion assemblies 60 and 64, as shown in FIGS. 3 and 4. When the flight mode is a cruising mode such as the horizontal flight, the tilting propulsion assemblies 58 may be tilted at a certain angle relative to the blade arrangement direction of the non-tilting propulsion assemblies 60 and 64, as shown in FIG. 5.

[0068] In some implementations, the air mobility apparatus 50 may include a gear 66 that is retractable or non-retractable for landing on land and / or water. The gear 66 may be located on both the front and rear of the air mobility apparatus 30, and it may include components such as wheels, treads, pontoons, or other components that help the air mobility apparatus land on land and / or water. Additionally, the air mobility apparatus 50 may have a stepper 68 that can be extended from the fuselage 52 for boarding and alighting, and side doors.

[0069] FIG. 6 is a diagram schematically illustrating a module constituting an air mobility apparatus. The following description focuses on the VTOL (Vertical Take-off and Landing) air mobility apparatus featuring motor-driven propulsion assemblies, as illustrated in FIGS. 2-5. However, the functionalities, modules, and apparatuses described In the present disclosure may also be applied to other types of air mobility apparatuses if they can be technically combined. Furthermore, the description of some of the functional modules in FIG. 6 is essentially the same as those in FIG. 1, so the corresponding descriptions will be brief.

[0070] The air mobility apparatus 100 may include a sensor unit 132, communication unit 134, manipulation unit 136, display 138, load device 140, power source unit 142, actuator 152, memory 154, and processor 156.

[0071] The sensor unit 132 may be equipped with various types of detectors that detect various states and situations arising in the internal and external environments of the air mobility apparatus 100. The sensor unit 132 may acquire sensor data used for flight control, status data that detects the status of the modules constituting the air mobility apparatus 100, and situational data that detects the situation of passengers and / or cargo. The sensor unit 132 may then provide this data to the processor 156, which processes the functions and operations.

[0072] Specifically, the sensor unit 132 may include a voltage / current sensor 132a, temperature sensor 132b, environmental sensor 132c, resistance sensor 132d, positioning sensor 132e, and a state recognition sensor 132f.

[0073] The voltage / current sensor 132a may, for example, detect the electrical status of at least one of the components, such as the electric battery of the power source unit 142, its charging module, or the motors and inverters of the propulsion assemblies 108a-108d, 110a-110d. The electric battery may be a secondary battery that can be charged externally but may not self-generate power. The charging module of the battery may be a self-generating battery, supplying power to the propulsion assemblies 108a-108d, 110a-110d. The self-generating battery may, for example, be a fuel cell module, though it is not limited to this. Based on the data, such as the electrical status and variations of the battery, motor, and inverter, detected by the voltage / current sensor 132, the processor 156 may detect flight modes, flight control status, battery control status, motor control status, inverter control status, and malfunction or abnormal states of these modules.

[0074] The temperature sensor 132b may detect the temperature of at least one component, such as the secondary battery of the power source unit 142, the self-generating battery, the motors and inverters of the propulsion assemblies 108a-108d, 110a-110d, or the cooling system that regulates the temperature of the propulsion assemblies 108a-108d, 110a-110d. In the case of the cooling system, the temperature sensor 132b may measure the temperature of the cooling medium circulating through the cooling system and estimate the temperature of the motors and inverters. Based on the temperature data detected by the temperature sensor 132b―such as the temperature of the battery, motors, and inverters―the processor 156 may detect flight modes, flight control status, battery control status, motor control status, inverter control status, and malfunction or abnormal situations of these modules.

[0075] The environmental sensor 132c may be equipped with, for example, cameras, radar sensors, LiDAR sensors, and other detectors to sense the external environment. The resistance sensor 132d may detect the operating resistance of at least one component, such as the secondary battery of the power source unit 142, the self-generating battery, or the motors and inverters of the propulsion assemblies 108a-108d, 110a-110d. Based on the resistance data detected by the resistance sensor 132d, the processor 156 may detect flight modes, flight control status, control status of the modules, and malfunction or abnormal situations of the modules.

[0076] The positioning sensor 132e is a module that identifies the location of the air mobility apparatus 100 and may include a GPS sensor or GNSS sensor, for example.

[0077] The state recognition sensor 132f may detect the motion corresponding to the operation of the aircraft, the load applied to the actuator 152, and the altitude of the aircraft. The state recognition sensor 132f may detect the three-axis status of the airframe 102, such as yaw, pitch, and roll, and output the motion status of the aircraft based on the aforementioned parameters. Such a state recognition sensor may, for example, include an IMU sensor or a gyro sensor. Additionally, the state recognition sensor 132f may be equipped with an altimeter to detect changes in the altitude of the aircraft. Furthermore, the state recognition sensor 132f may include a load detection sensor that measures the load applied to at least one of the propulsion assemblies 108a-108d, 110a-110d , the main wing 104, or the tail wing 106. The load detection sensor may be provided in a specific part of the airframe. The aircraft load may refer to the force or pressure exerted by surrounding airflows on at least one of the actuators 152, wings 104, 106, or the airframe 102 during flight. The state recognition sensor 132f may include a barometric pressure sensor and an altitude sensor. The barometric pressure sensor may detect the atmospheric pressure around the aircraft or the connection pressure between the air mobility apparatus 100 and an external device connected to it. The altitude sensor may be a sensor that identifies the current altitude of the aircraft.

[0078] The above descriptions of the data from the sensor unit 132 are merely illustrative, and the sensor unit may also include additional sensor data that detects various situations not listed here.

[0079] The manipulation unit 136 may include a module that receives flight control input from the pilot. For example, the manipulation unit 136 may include at least one inceptor inside the airframe. In addition to the hardware components, the manipulation unit 136 may further include user interfaces such as soft keys. The user interface may be implemented on the display 138, for example. The display 138 may be controlled by the processor 156 to show the attitude status of the air mobility apparatus 100, control status, route information, remaining energy level, and the surrounding environment as detected by the environmental sensor 132c. While the present disclosure provides for the manipulation unit 136 for manual piloting, autonomous flight may also be executed upon the pilot's request.

[0080] The power source unit 142 may generate and supply power to the propulsion assemblies 108a-108d, 110a-110d, and the load device 140. Referring to FIG. 6, the power source unit 142 may be configured to include three battery modules. The power source unit 142 may include a first battery, a second battery and a third battery.

[0081] The first battery, for example, may be a self-generating battery unit such as a fuel cell module 146. The first battery may be configured to generate power and include a battery module or at least one battery configured to store the generated power. For example, the fuel cell module 146 may include a fuel cell stack, a fuel supply unit, an air supply unit, a water treatment unit, a power conversion unit, and a control unit. In some examples, where the battery has self-generating capabilities. The first battery is not limited to a fuel cell. For instance, in some examples, the first battery may include a power generator, solar panel system, engine, etc. The fuel cell module 146 may supply power to at least one of the propulsion assemblies 108a-108d, 110a-110d, the second battery module, and the third battery module, depending on the flight mode, flight situation, and flight route.

[0082] The second and third batteries may contain different types of secondary batteries. Depending on the flight mode, flight situation, and flight route, the second and third batteries may supply power to the propulsion assemblies 108a-108d, 110a-110d, or be recharged by the fuel cell module 146. The second and third batteries may include the high-power battery module 148 and the high-energy battery module 150, respectively. Each battery module 148, 150 does not refer to a specific component of a battery, and may include a battery cell, cell assembly and battery pack. The high-power battery module 148 may have a higher discharge characteristics than the high-energy battery module 150 and may also be referred to as a high-discharge battery module.

[0083] The high-power battery module 148 may be a secondary battery charged and discharged at a C rate of 5C or greater and less than 10C. The high-energy battery module 150 may be a secondary battery with a higher energy capacity than the high-power battery module 148 but lower instantaneous output characteristics and may have a C rate of 0.5 to 2.

[0084] The high-power battery module 148 and the high-energy battery module 150 may be lithium-based batteries with either different compositions or the same composition with different ratios. More specifically, the high-power battery module 148 and the high-energy battery module 150 may include ternary materials in NCM (Nickel-Cobalt-Manganese) or NCA (Nickel-Cobalt-Aluminum) with different ratios. In FIG. 6, the high-power battery module 148 and the high-energy battery module 150 are illustrated in combination with the fuel cell module 146. However, in other implementations related to FIGS. 19-21, only one of these batteries may be combined with the fuel cell module 146. The present disclosure will first describe an implementation involving three batteries.

[0085] The actuator 152 may include multiple propulsion assemblies 108a-108d, 110a-110d and may also be referred to as the actuating unit. The multiple propulsion assemblies 108a-108d, 110a-110d may be the components corresponding to the driving unit 22 of FIG. 1in the air mobility apparatus 100.

[0086] The propulsion assemblies 108a-108d, 110a-110d may include propellers with multiple blades, hubs, motors, inverters, and propulsion cooling modules. As shown in FIG. 7, the present disclosure illustrates that the air mobility apparatus 100 features a tilting rotor type. In the case of tilting rotors, at least some of the multiple propulsion assemblies 108a-108d, 110a-110d may be able to tilt. As illustrated in FIG. 7, the first propulsion assemblies 108a-108d are configured as tiltable rotors, and the second propulsion assemblies 110a-110d are configured as fixed rotors. Here, fixed rotors refer to rotors that do not tilt during flight and maintain a fixed blade arrangement.

[0087] The first propulsion assemblies 108a-108d may be installed on the main wing 104 and the tail wing 106 and may be hinged to booms extending from the wings. The first propulsion assemblies 108a-108d may be equipped with nacelles 116 that house motors 158 that power the propellers 112, and the nacelles 116 may tilt relative to the booms 118. The positions and tilting structures of the first propulsion assemblies 108a-108d are not limited to the example in FIG. 7 and may be configured in various ways. The first propulsion assemblies 108a-108d may operate in hovering mode, cruising mode and transition phase between these modes. In hovering mode, the first propulsion assemblies 108a-108d may be configured with the blade arrangement shown in FIG. 7, and in cruising mode, the first propulsion assemblies 108a-108d may be operated to have the same blade arrangement as the second propulsion assemblies 110a-110d shown in FIG. 7. The second propulsion assemblies 110a-110d may serve as lift rotors and be fixedly installed on the booms extending from the main wing 104. The positions and tilting structures of the second propulsion assemblies 110a-110d are not limited to the example in FIG. 7 and may also be configured in various ways. The second propulsion assemblies 110a-110d are mainly used in hovering mode, but they may also operate depending on the flight situation during horizontal flight. In the present disclosure, the first propulsion assemblies 108a-108d and the second propulsion assemblies 110a-110d may each be referred to as a single propulsion assembly or a propulsion assembly set. For example, a set of the first propulsion assembly with multiple tilting rotors may be collectively referred to as the first propulsion assembly 108a-108d, and a set of the second propulsion assembly with multiple fixed rotors may be collectively referred to as the second propulsion assembly 110a-110d.

[0088] As another example than Fig. 7, the tilting rotor type may be a type in which all of the multiple propulsion units function as tilting rotors. The implementation of the present disclosure may be applied to both lift-and-cruise types and multicopter types.

[0089] In some examples, The power source unit 142 may be equipped with a single module energy source for the first and second propulsion units 108a to 108d, 110a to 110d, and as another example, the power source unit 142 may include energy sources of multiple modules for stability, redundancy, and stable thrust of energy supply to the first and second propulsion units 108a to 108d, 110a to 110d. Additionally, the multiple propulsion assemblies 108a-108d, 110a-110d may be electrically connected to at least one of the multiple energy sources for the purposes described above. FIGS. 8 and 9 provide examples of layouts showing the energy supply architecture between the multiple modules of the power source unit and the propulsion assemblies.

[0090] FIG. 8 is a diagram illustrating an example of the layout of the energy supply architecture between the power source unit and the propulsion assembly.

[0091] In FIG. 8, the first to third energy sources 142a-142c constituting the power source unit 142 may each include one of the following: electric batteries of the same performance, electric batteries with heterogeneous characteristics, a combination of an electric battery and another type of power source, or different types of power sources other than electric batteries. Electric batteries with the same characteristics may be, for example, the high-power battery module 148. Electric batteries with different characteristics may be the high-power battery module 148 and the high-energy battery module 150. Other types of power sources may be, for example, a fuel cell module 146 or an internal combustion engine coupled with a generator.

[0092] Other types of power sources may be, for example, self-generating power sources or, power sources that do not have self-generation capabilities. Other types of power sources are not limited to the examples mentioned and may include any source capable of supplying power to the propulsion assemblies 146a-146d in a configuration different from electric batteries.

[0093] In FIG. 8, each energy source 142a-142c is electrically connected to at least two propulsion assemblies 152a-152d. The propulsion assemblies 152a-152d depicted in FIG. 8 may respectively correspond to the first and second propulsion assemblies 108a-108d, 110a-110d illustrated in FIG. 6, depending on the design. The energy sources 142a-142c and propulsion assemblies 152a-152d may be electrically connected in a loop configuration. This allows the air mobility apparatus 100 perform a redundancy plan to stably supply power to the propulsion assemblies 152a to 152d using other energy sources 142a, 142c in the event of an abnormality in a specific energy source 142b.

[0094] In addition, in the event of a failure such as an abnormal decrease in the thrust of a specific propulsion assembly 152b, it may be considered that sufficient power is supplied to other propulsion assemblies in order to increase the thrust of other propulsion units that are not in a predetermined corresponding relationship with the specific propulsion assembly 152b. Also, this is to ensure a stable supply of power to propulsion assemblies that require increased power, to increase the power for some of the propulsion assemblies 152a to 152d depending on the flight mode.

[0095] FIG. 9 is a diagram illustrating another example of the layout of the energy supply architecture between the power source unit and the propulsion assembly.

[0096] In FIG. 9, the power source unit 142 is shown to include a first energy source 142d and a second energy source 142e. The first energy source 142d supplies power to the propulsion assemblies 146e to 146h regardless of the flight mode, while the second energy source 142e supplies supplemental power to at least some of the propulsion assemblies 152e to 152h along with the first energy source 142d in flight modes with high power consumption, such as hovering mode.

[0097] The first energy source 142d may include at least one energy source to supply power to the corresponding propulsion assemblies 152e to 152h, as described in FIG. 9. The second energy source 142e may comprise at least one of electric batteries with the same performance as the first energy source 142d, electric batteries with different performance, or another type of power source.

[0098] In FIG. 9, in a flight mode with relatively low power consumption, such as cruising mode, the first energy source 142d may supply power to the propulsion assemblies 152e to 152h. Additionally, instead of supporting the power supply to the propulsion assemblies 152e to 152h, the second energy source 142e may charge the first energy source 142d based on the energy remaining, generation performance, current, temperature, and other conditions of the first energy source 142d and the second energy source 142e. In another example, the first energy source 142d may charge the second energy source 142e based on the parameters described above. Moreover, the second energy source 142e may also provide supplemental power to the propulsion assemblies 152e to 152h depending on the situation, even in flight modes with low power consumption. As described above, in flight modes with high power consumption, the second energy source 142e may supply power to at least some of the propulsion assemblies 152e to 152h along with the first energy source 142d based on the energy remaining, generation performance, current, temperature, and other conditions of the energy sources. In another example, if the power supply of the first energy source 142d decreases due to its condition, the second energy source 142e may provide power to the corresponding propulsion assemblies 152e to 152h based on the incremental power required by the mode and the decreased power of the first energy source 142d. In Fig. 8, although the first and second energy sources, 142d and 142e, are shown as single sources, each of the first energy source 142d and the second energy source 142e may be provided in plurality, depending on design specifications that include the performance of the propulsion assembly, the number of propulsion assemblies, the installation position of the propulsion assemblies, and the required output power.

[0099] As another example different from the above description, Fig. 9 may be understood as an extension of the circular-type redundancy shown in Fig. 8. The first energy source 142d and the second energy source 142e may be connected to each of the propulsion assemblies 146e - 146h. Specifically, the third energy source 142c cannot support the first and second propulsion assemblies 146a and 146b in the redundancy of Fig. 8, however each energy source 142d and 142e may be connected to all propulsion assemblies 152e through 152h in the redundancy of Fig. 9. According to the redundancy in Fig. 9, each energy source may primarily support at least one propulsion assembly under normal situations, and in an emergency, each energy source may supply energy to at least one other propulsion assembly according to a redundancy plan. Fig. 9 illustratively depicts two energy sources 142d and 142e. However, the present disclosure is not limited thereto, and three or more energy sources may be provided. For example, the energy sources may be provided in three or more in accordance with design specifications including the performance of the propulsion assembly, the number of propulsion assemblies, the installation position of the propulsion assemblies, and the required output power and so on.

[0100] Returning to FIG. 6, the memory 154 stores applications and various data for flight control of the air mobility apparatus 100, and upon request from the processor 156, it can load applications or read and write data. In the present disclosure, the applications and data may relate to the control of multiple power sources depending on the operational state of the air mobility apparatus 100. The data may include sensor data, the state of the air mobility apparatus, and flight control parameters. Additionally, the applications and data may store and manage data and software mentioned in FIG. 1.

[0101] In the present disclosure, the processor 156 uses the applications, instructions, and data stored in the memory 154 to process tasks such as flight control, route control, power and energy control via the powertrain, and autonomous movement control. When the power source unit 142 includes three battery modules operating in a normal state as in FIG. 6, the processor 156 may control the power source unit 142 to supply the power output from both the fuel cell module 146 and the high-power battery module 148 to the first propulsion assemblies 108a-108d in response to high-power mode of the actuators 152, and to supply power output from both the high-energy battery module 150 and the high-power battery module 148 to the second propulsion assemblies 110a-110d. In response to a low-power mode of the actuators 152, the processor 156 may control the power source unit 142 to supply the power output from the fuel cell module 146 to at least the first propulsion assemblies 108a-108d.

[0102] The high-power mode may refer to a mode that uses more power than a baseline power to execute flight operations. Details of the baseline power will be described later. The high-power mode may be applied, for example, to hovering mode performed in takeoff and landing or to specific flight operations occurring during cruising. Hovering mode exceeding the baseline power is not limited to takeoff and landing but may be activated in situations requiring hovering operation in accordance with the flight status. Cruising flight refers to operations in cruising mode. Specific operations during cruising flight may require flight power exceeding the baseline power. Specific operations may include, for example, rapid ascent, sharp turns, responding to headwinds or jet streams, and increasing thrust due to deteriorating weather conditions. These operations may require substantial power for at least one of the first and second propulsion assemblies 108a-108d, 110a-110d.

[0103] Whether high-power mode is entered may be determined based on at least one of flight control commands and sensor data. Flight control commands may be or include control request signals related to high-power mode issued by the pilot or processor 156. The flight control commands may be, for example, request signals related to takeoff, landing, or specific operational controls. Sensor data may include data from voltage / current sensors 132a, temperature sensors 132b, and / or resistance sensors 132d, which detect the state of at least one of the motor and inverter. Sensor data may also include changes in flight attitude and altitude detected continuously by the state recognition sensor 132f. Additionally, when the processor 156 performs speed control to reduce the difference between the pilot's speed command and the actual aircraft speed, whether high-power mode is entered may be determined based on current sensing data from the inverters of the propulsion assemblies 108a-108d, 110a-110d and / or current data measured by the converters (e.g., FDC 166 and BHDC 170 in FIG. 10) connected to the battery modules 146-150.

[0104] For instance, the low-power mode may be a mode applied to operations of the air mobility apparatus related to cruising flight. Cruising flight refers to flight occurring between takeoff and landing, typically involving horizontal flight at a predetermined altitude range near the destination based on the flight plan and airspace. Low-power mode may require flight power less than the baseline power.

[0105] The low-power mode may also be determined based on at least one of flight control commands and sensor data. Flight control commands may be or include control request signals related to low-power mode issued by the pilot or processor 156. The flight control commands may be, for example, request signals related to cruising flight or horizontal flight. Additionally, sensor data, as described earlier, may include the state of at least one of the motor, inverter, and converter detected by the sensor unit 132. If the voltage / current sensor 132a, temperature sensor 132b, and resistance sensor 132d output a state defined differently from high-power mode, low-power mode may be determined. If changes in flight attitude and altitude detected by the state recognition sensor 132f continue within a predetermined range, low-power mode may be deemed active.

[0106] In some examples, when the power source unit 142 includes two battery modules operating in a normal state, as shown in FIG. 21, the processor 156 may control the power source unit 142 to supply power output respectively from the fuel cell module 246 and the high-power battery module 250a to the first propulsion assemblies 108a-108d in response to high-power mode. Additionally, in high-power mode, the processor 156 may control the power source unit 142 to supply power output respectively from the fuel cell module 246 and the high-power battery module 250b to the second propulsion assemblies 110a-110d. In low-power mode with two battery modules, the processor 156 may control the power source unit 142 similarly to the process described with three battery modules.

[0107] In relation to the present disclosure, the processor 156 may control the power distribution to each propulsion assembly from the battery modules via redundancy when portion of the multiple battery modules are in a failure state. The detailed aspects of the redundancy control will be described later. For example, the failure state may refer to a state in which the battery module provides no power or power less than a target power range corresponding to a specification of the battery module. In some cases, a normal state may refer to a state in which the battery module provides power within the target power range.

[0108] The processor 156 may include a single processing unit or include multiple processing units for distributed processing. Multiple processing units may include, for example, an APS (Air Processing System), FCC (Flight Control Unit), PCU (Power Control Unit), battery controllers, TMS (Thermal Management System), FPS (Fuel Processing System), and others, but they are not limited to these examples.

[0109] In some examples, even if the operations according to the present disclosure are distributed among multiple systems or controllers, for convenience of explanation, they may collectively be referred to as the processor 156. The processor 156 is described as handling the functions, operations, and processes according to this implementation.

[0110] The present disclosure primarily describes a VTOL-type air mobility apparatus 100. In some examples, where other types of mobility apparatuses, such as ground mobility apparatuses, robots, drones, or others, are equipped with similar systems, the technical concepts of the present disclosure and the implementations may also be applied to those mobility apparatuses.

[0111] Hereinafter, referring to FIGS. 10 to 16, the description of the air mobility apparatus and the power control operations in the multiple power sources will be provided. FIG. 10 is a diagram illustrating an example of the connection between the power source unit and the propulsion assembly. FIG. 10 primarily illustrates the components related to the power source unit 142 and the propulsion assemblies 108a, 110a from FIG. 6, with the power source unit 142 having three battery modules 146-150. Although FIG. 10 shows one example where the first propulsion assembly 108a and the second propulsion assembly 110a are each singular, the implementation in FIGS. 10 to 16 can be commonly applied to other propulsion assemblies 108b-108d and 110b-110d. The high-power battery module 148 and the high-energy battery module 150 are not limited to being singular but can further comprise multiple battery modules connected to other propulsion assemblies 108b-108d and 110b-110d.

[0112] The fuel cell module 146 and the high-power battery module 148 can be connected in parallel to a common node 1. The fuel cell module 146 can be selectively connected to the high-power battery module 148 to supply charging power to it. The selective connection may be implemented by a switching element of the high-power battery module 148. The fuel cell module 146 is electrically connected to the first propulsion assembly 108a, while the high-power battery module 148 can be selectively connected to the first and second propulsion assemblies 108a and 110a. This selective connection can also be executed by the switching element of the high-power battery module 148. The high-power battery module 148 may include a protection element 174 that disconnects the electrical connection with the first and second propulsion assemblies 108a and 110 in response to a failure state. The protection device 176 may be an assembly or a switching device that includes a fuse. The assembly may be, for example, a PreCharge-Relay-Assembly (PRA) with a pyro fuse that physically disconnects the circuit when an overcurrent or overvoltage occurs. The fuel cell module 146 includes a control device to efficiently manage and distribute power in the fuel cell system, and this control device may be referred to as the PMC (Power Management Controller).

[0113] As shown in FIGS. 10 and 11, a fuel cell converter 166 and a Reverse current prevention device 168 can be arranged between the fuel cell module 146 and node 1 (marked as #1 in FIG. 11). FIG. 11 is a diagram showing an example of the circuit of the power source unit. The fuel cell converter 166 may be configured as a unidirectional converter. The fuel cell converter 166 regulates the output distribution between the fuel cell module 146, the high-power battery module 148, and the high-energy battery module 150, and may control the input voltage to prevent reverse current flowing into the PMC. The fuel cell converter 166 may be an FDC (Fuel Cell DC / DC Converter). The Reverse current prevention device 168 allows the current output from the fuel cell module 146 to flow to node 1 while preventing reverse current from node 1 back to the fuel cell module 146. The Reverse current prevention device 168 may be, for example, a unidirectional diode. The FC BOP may be a unit used to return the output current of the fuel cell module 146 to the fuel cell module 146.

[0114] As shown in FIGS. 10 and 11, the high-power battery module 148 and the high-energy battery module 150 can be connected in parallel to the bidirectional converter 170 between node 1 and node 2. The high-energy battery module 150 can be selectively connected to the second propulsion assembly 110a and the fuel cell module 146. This selective connection can be implemented by a switching element of the high-energy battery module 150.

[0115] The bidirectional converter 170 may be arranged between the high-power battery module 148 and the high-energy battery module 150. The bidirectional converter 170 can be a converting device configured with a half-bridge or H-bridge structure, enabling charging and discharging functions for the battery modules and power transmission characteristics. The bidirectional converter 170 may be, for example, a BHDC (Bi-directional High Voltage DC / DC Converter). The BHDC converter may include an input voltage controller, output voltage controller, and a current controller for the inverter. The output voltage controller manages the voltage supplied to the inverter, while the input voltage controller regulates the voltage applied from the fuel cell module 146. The current controller may be controlled by limit logic for current command. The limit logic for current command may include, for example, external current commands, current limits based on the input voltage, and current limits due to the temperature of the inverter. The bidirectional converter 170 illustrated in Figure 11 may include a high-side switch 178 that controls switching on the high-voltage side and a low-side switch 180 that controls switching on the low-voltage side.

[0116] The high-energy battery module 150, in response to a failure state, may include a safety element 176 that disconnects the electrical connection with the first and second propulsion assemblies 108a and 110. The safety element 176, similar to the protection element 174, may be an assembly with a fuse or a switching device.

[0117] As illustrated in FIG. 12, the first and second propulsion assemblies 108a-108d and 110a-110d may each include first and second motors 158, 162 and first and second inverters 160, 164. FIG. 12 is a diagram showing an example of the circuit of the propulsion assembly. Each motor 158, 162 may be configured as a dual 3-phase motor for redundancy, and the inverters may also be configured as dual inverters. Each inverter 160, 164 can apply alternating current to the respective motor 158, 162. For example, the inverters 160, 164 can determine the amount of current supplied to the motors 158, 162 based on the flight mode, the condition of the motor / inverter, the attitude of the aircraft, and the load condition of the aircraft. Flight modes may include cruising mode and hovering mode. The condition of the motor / inverter may include, for example, temperature, voltage, current, and resistance of motor / inverter detected via the sensor unit 132. The attitude of aircraft may include longitudinal / lateral posture information of aircraft recognized by the state recognition sensor 132f. The load status of aircraft may refer to the load weight caused by passengers and cargo.

[0118] FIG. 13 is a diagram showing an example of the profile of the first propulsion assembly by flight mode. The flight modes are categorized into the takeoff phase, cruising phase, and landing phase. Here, the takeoff and landing phases may be the phases in which the first propulsion assembly 108a operates in hovering mode. The hovering mode may be a mode in which the first and second propulsion assemblies 108a and 110a operate to generate lift for the air mobility apparatus 100. The hovering mode may be a type of high-power mode. The cruising phase may be a mode in which at least the first propulsion assembly 108a operates to generate thrust. The second propulsion assembly 110a may also operate during the cruising phase to ensure stable flight operations (or flight attitude). The cruising phase may be a type of low-power mode. The transition phase may refer to the phase corresponding to the transition of flight operation that occurs between the takeoff / landing phases and the cruising phase. In the present disclosure, the transition phase may be considered part of the takeoff / landing phase or part of the cruising phase. The first propulsion assembly 108a may perform tilting operations during the transition phase. The second propulsion assembly 110a may operate during the transition phase to maintain appropriate lift and stable flight operations.

[0119] The takeoff and landing phases may require high power since they involve generating lift by inducing airflow through the rotation of the rotor. The portions enclosed by the profile for each phase illustrated in FIG. 13 represent the electric power or energy applied during those phases.

[0120] For instance, the required output for the first propulsion assembly 108a during the takeoff phase may be determined based on the average value of the accumulated takeoff power during the flight and the takeoff power determined based on the takeoff environment and the flight operation status. The required output for the takeoff phase may be the power allocated to the first propulsion assembly 108a based on at least one of the average value and the takeoff power. Specifically, the takeoff environment includes conditions from the takeoff point to the cruising point, such as airflow, weather, and obstacles location along the path, but is not limited to these factors. The flight operation status includes factors such as the load weight, the in-cabin arrangement of cargo, the time of day (day or night), the entire flight path, and the flight duration, but is not limited to these.

[0121] In FIG. 13, the required output of the first propulsion assembly 108a is shown to be uniform during the takeoff phase. However, during actual flight, the required power from the beginning of takeoff to the transition phase may have maximum instantaneous power, and the power from the transition phase to the cruising phase may be lower than the maximum instantaneous power. The maximum required output of the first propulsion assembly 108a may be determined based on at least one of the maximum instantaneous power, the average value, and the takeoff power. Accordingly, in another example, the required output for the takeoff phase may be determined by the maximum required output.

[0122] The required output for the takeoff phase may be defined as the sum of detailed outputs. Specifically, the required output for the takeoff phase may be the sum of the baseline power and the increment of the maximum required output. The baseline power may be set based on the power applied to the first propulsion assembly in the low-power mode. For example, the baseline power may be set based on the power consumed by the first propulsion assembly 108a during the cruising phase, which will be explained later. The baseline power may, for example, be the power consumed by the first propulsion assembly 108a during the cruising phase, or in another example, may include the power required for supporting the first propulsion assembly 108a and the battery modules 148, 150 (e.g., charging power). Detailed information about the supporting power will be explained later. The increment may refer to the additional power required by the first propulsion assembly 108a only during the takeoff phase. The increment may, for example, be determined based on the difference between the maximum required output and the baseline power.

[0123] The required output of the first propulsion assembly 108a during the cruising phase may be determined based on the baseline power. The baseline power may be set to provide a consistent output of power. The baseline power may be based, for example, on the average accumulated cruising power during the flight or cruising power determined based on various data. The required output for the cruising phase may be the power allocated to the first propulsion assembly 108a based on at least one of the average value and the cruising power. The cruising power may be generated based on the flight environment and the flight operation status. The flight environment includes various conditions during the cruising phase along the route, such as airflow, weather, and obstacles location along the path, but is not limited to these factors. The flight operation status includes factors such as load weight, the in-cabin arrangement of cargo, the time of day (day or night), the entire flight path, and the flight duration, but is not limited to these.

[0124] The fuel cell module 146 can charge at least one of the high-power battery module 148 and the high-energy battery module 150 during flight in the cruising phase based on the state of charges of the high-power battery module 148 and the high-energy battery module 150. In another example, the baseline power may further include supporting power determined based on the states of the battery modules 148 and 150. Specifically, the baseline power may be determined based on the average value, the cruising power, and the supporting power. The states of the battery modules 148 and 150 may include the charging status, temperature, and operational status of each battery module 148 and 150. The supporting power based on the charging status of the battery may be referred to as charging power.

[0125] The cruising phase may include an inertial flight phase before the landing phase. When the aircraft reaches the vicinity of the destination, the pilot may generate a request to gradually reduce the speed of the aircraft. The processor 156, such as the FCC, may receive this request and send a reduced current command to the first propulsion assembly 108a. Due to the inertia generated by the speed of the aircraft prior to the deceleration request, the inertial flight phase may be a phase operated with less energy than other ranges of the cruising phase. To charge the high-power battery module 148 and the high-energy battery module 150, the fuel cell module 146 may utilize the significant energy not delivered to the first propulsion assembly 108a during the inertial flight phase. The unutilized energy can be determined by subtracting the power required for the first propulsion assembly 108a during the inertial flight phase from the power of the fuel cell module 146 during the cruising phase. For example, the PMC can charge the aforementioned battery modules by transmitting negative (-) current through the BHDC 170 based on the state of charge.

[0126] The required output of the first propulsion assembly 108a during the landing phase may, for example, be substantially the same as in the takeoff phase.

[0127] In another example, the required output of the landing phase may differ from the takeoff phase. The required output for the landing phase may be based, for example, on the average value of the accumulated landing power during the flight, the landing environment, and the flight operation status. The required output for the landing phase may be the power allocated to the first propulsion assembly 108a based on at least one of the average value and the landing power. Specifically, the landing environment includes the conditions from the inertial flight point (or the corresponding point of the transition phase) to the landing point, such as airflow, weather, and obstacle location along the path, but is not limited to these. The flight operation status may be the same as or different from the takeoff phase. The flight operation status includes factors such as load weight, the in-cabin arrangement of cargo, the time of day (day or night), the entire flight path, and the flight duration, but is not limited to these.

[0128] In FIG. 13, the required output of the first propulsion assembly 108a during the landing phase is shown as uniform, but in actual flight, the required power may have maximum instantaneous power from the timing before landing until the transition phase. The maximum instantaneous power during the landing phase may be the same as or different from that of the takeoff phase. The maximum required output during the landing phase may, for example, be the same as the maximum instantaneous power of the takeoff phase. The maximum required output may, in another example, be determined based on at least one of the maximum instantaneous power identified during the landing phase, the average value, and the landing power. Furthermore, the required output of the takeoff phase may be determined based on the maximum required output.

[0129] The required output for the landing phase may be defined as the sum of detailed outputs. Specifically, the required output for the landing phase may be the sum of the baseline power and the increment of the maximum required output. The baseline power and the increment are substantially the same as described for the takeoff phase.

[0130] FIG. 14 is a diagram showing an example of the profile of the second propulsion assembly by flight mode.

[0131] The flight modes are categorized into the takeoff phase, cruising phase, and landing phase, the same as in FIG. 13. Furthermore, the transition phase may belong between the takeoff / landing phases and the cruising phase. The takeoff and landing phases may be phases in which the second propulsion assembly 110a operates in hovering mode. The cruising and transition phases are the same as described in FIG. 13, and the areas enclosed by the profile for each phase illustrated in FIG. 14 are also the same as those mentioned in FIG. 13 .

[0132] The required output of the second propulsion assembly 110a during the take-off phase may, for example, be the same as that of the first propulsion assembly 108a during the take-off phase. The required output of the second propulsion assembly 110a during the take-off phase can be determined based on the data described in Figure 13. This data may be the take-off power generated based on the average value of the accumulated take-off power during flight, the take-off environment, and the flight operation status. The required output during the take-off phase may be the power assigned to the second propulsion assembly 110a from at least one of the average value and the take-off power. The required output of the second propulsion assembly 110a may, in other cases, be different from that of the first propulsion assembly 108a during the take-off phase.

[0133] In Figure 14, the required output of the second propulsion assembly 110a is shown to be uniform during the take-off phase, but as explained in Figure 13, in actual flight, the required power may have maximum instantaneous power during a certain range of the take-off phase. The maximum required output of the second propulsion assembly 110a can be determined based on at least one of the maximum instantaneous power, the average value, and the take-off power. Accordingly, the required output during the take-off phase may be determined by the maximum required output in other cases.

[0134] As explained in Figure 13, the required output of the second propulsion assembly 110a during the take-off phase can be defined as the sum of detailed power outputs. Specifically, the required output during the take-off phase may be the sum of the baseline power and the increment of the maximum required output. The baseline power of the second propulsion assembly 110a may be, for example, the power consumed by the second propulsion assembly 110a during the cruising phase, or in other cases, it may include the power of the second propulsion assembly 110a and the supporting power for the battery modules 148 and 150. In the present disclosure, as illustrated in Figures 13 and 14, the baseline power of the second propulsion assembly 110a may be the same as that of the first propulsion assembly 108a. The increment may be the additional power required by the second propulsion assembly 110a only during the take-off phase. The increment may be determined, for example, based on the difference between the maximum required output and the baseline power. If the maximum required output of the second propulsion assembly 110a is different from that of the first propulsion assembly 108a, the increment may also differ from the increment of the first propulsion assembly 108a.

[0135] Since the second propulsion assembly 110a mainly operates to generate lift, it may not be operational during the cruising phase. For stable flight operation (or flight attitude), the second propulsion assembly 110a may be operated during the cruising phase. The required output of the second propulsion assembly 110a during the cruising phase may be generated based on a current command related to the control request from the pilot or processor 156. When the second propulsion assembly 110a is operated during the cruising phase, the fuel cell module 146 may supply power to the second propulsion assembly 110a according to the required output. During the power supply to the second propulsion assembly 110a, the fuel cell module 146 may charge at least one of the high-power battery module 148 and the high-energy battery module 150 based on the state of charge of the battery modules.

[0136] The required output of the second propulsion assembly 110a during the landing phase may, for example, be substantially the same as during the take-off phase. In another example, the required output during the landing phase may be different from that of the take-off phase. The required output during the landing phase can be determined based on the data described in Figure 13. This data may be the landing power generated based on the average value of the accumulated landing power during the flight, the landing environment, and the flight operation status. The required output during the landing phase may be the power assigned to the second propulsion assembly 110a from at least one of the average value and the landing power.

[0137] In Figure 14, the required output of the second propulsion assembly 110a is shown to be uniform during the landing phase, but as explained in Figure 13, in actual flight, the required power may have maximum instantaneous power during a certain range of the landing phase. The maximum required output of the second propulsion assembly 110a may be determined based on at least one of the maximum instantaneous power, the average value, and the landing power. Accordingly, the required output during the landing phase may be determined by the maximum required output in other cases.

[0138] The required output of the second propulsion assembly 110a during the landing phase may be the sum of the baseline power and the increment of the maximum required output. The baseline power and the increment are substantially the same as described for the take-off phase.

[0139] Figures 13 and 14 describe the high-power mode in terms of hovering flight during take-off and landing, but the high-power mode may refer to a mode in which the flight operation is carried out using more power than the baseline power. The high-power mode may be applied not only to hovering flight but also to specific flight operations during the cruising phase. The details of specific operations are the same as described in Figure 6. The output profile of the first and second propulsion assemblies 108a and 110a according to the high-power mode related to specific operations may differ from those shown in Figures 13 and 14. The required output and increment shown in Figures 13 and 14 may be set according to specific operations. Whether the high-power mode or low-power mode is activated may be determined based on at least one of the flight control command and the sensor data. The details of the aforementioned are the same as described in Figure 6.

[0140] FIG. 15 is a diagram showing power supply to the propulsion assembly in high-power mode with three multiple power sources operating in a normal state. Here, the hovering mode is exemplified as the flight mode applied during the take-off or landing phase shown in Figures 13 and 14. Since the hovering mode is a type of high-power mode, it may be described interchangeably with the high-power mode below. Figure 15 shows the power control of the battery in high-power mode. Figures 15 and 16 illustrate one first propulsion assembly 108a and one second propulsion assembly 110a, respectively. However, the following description related to Figures 15 and 16 can also apply to other first propulsion assemblies 108b-108d and other second propulsion assemblies 110b-110d.

[0141] Processor 156 may determine whether the propulsion assembly has entered the hovering mode (or high-power mode) based on at least one of the flight control command and sensor data. In the present disclosure, the flight control command is a control request signal from the pilot or processor 156, for example, a request signal related to take-off, landing, or specific operation control. The sensor data may be output from the sensor unit 132 which detects state in at least one of the motor, inverter, flight attitude change, and altitude change.

[0142] For hovering related to take-off, processor 156 may tilt the first propulsion assembly 108a into a hovering arrangement and drive the second propulsion assembly 110a based on the control request signal. When the first propulsion assembly 108a is tilted into a hovering arrangement, processor 156, including the FCC and PCU, may generate a current increase command based on a specific output command and transmit it to the fuel cell converter FDC 166 and the bidirectional converter 170. The fuel cell converter 166 and the bidirectional converter 170 may control the current to be applied to the first and second propulsion assemblies 108a and 110a based on the current increase command. Processor 156 may determine that the flight mode has entered the high-power mode in response to the current increase command, applied current, and tilt control. In the transition phase where the air mobility apparatus 100 approaches the target altitude, processor 156 may gradually tilt the first propulsion assembly 108a into a cruise arrangement and gradually reduce the rotation of the second propulsion assembly 110a. If the required outputs of the first and second propulsion assemblies 108a and 110a exceed the baseline power during the transition phase, processor 156 may maintain the high-power mode.

[0143] For hovering related to landing, when the air mobility apparatus 100 approaches the destination, processor 156 may reduce the horizontal speed to zero based on the control request signal. Additionally, processor 156 may tilt the first propulsion assembly 108a into a hovering arrangement and drive the second propulsion assembly 110a based on the control request signal. Processor 156 may generate a maximum output command to compensate for the reduced lift according to the decreasing horizontal speed.

[0144] For example, FCC may generate a maximum output command and transmit it to PCU, which, based on the maximum output command, may generate a maximum current command and transmit it to the fuel cell converter FDC 166 and the bidirectional converter 170. The fuel cell converter 166 and the bidirectional converter 170 may control the current to be applied to the first and second propulsion assemblies 108a and 110a based on the current increase command. Processor 156 may determine that the flight mode has entered the high-power mode in response to the maximum output command, maximum current command, applied current, and tilt control. In the transition phase where the air mobility apparatus 100 descends from cruise flight, processor 156 may gradually tilt the first propulsion assembly 108a into a hovering arrangement and gradually increase the rotation of the second propulsion assembly 110a. If the required outputs of the first and second propulsion assemblies 108a and 110a exceed the baseline power during the transition phase, processor 156 may maintain the high-power mode.

[0145] Apart from the examples illustrated above, the details related to entering the high-power mode are described in Figure 6 and will be omitted here.

[0146] Upon detecting the hovering mode (i.e., the high-power mode) due to take-off or landing, processor 156 may supply the power output respectively from the fuel cell module 146 and the high-power battery module 148 to the first propulsion assembly 108a. Processor 156 may determine the power output of the fuel cell module 146 and the high-power battery module 148.

[0147] The power of the fuel cell module 146 is determined based on the baseline power of the first propulsion assembly 108a, allowing the fuel cell module 146 to support the power required by the first inverter 160 in high-power mode. The baseline power may be set based on the power applied to the first propulsion assembly 108a in low-power mode. Processor 156 may determine the power of the fuel cell module 146 to be below the baseline power based on the flight state and the condition of each battery module. In the present disclosure, the fuel cell module 146 is illustrated as supplying baseline power to the first propulsion assembly 108a, as shown in Figure 13. Therefore, the power supplied by the fuel cell module 146 to the first propulsion assembly 108a in both high-power mode and low-power mode may be the same. The detailed explanation of baseline power is provided in Figure 13 and will be omitted here.

[0148] The power supplied by the high-power battery module 148 to the first propulsion assembly 108a may be determined based on the difference between the required output of the first propulsion assembly 108a in high-power mode and the baseline power of the first propulsion assembly 108a. Here, the required output of the first propulsion assembly 108a may be set as the maximum required output of the first propulsion assembly 108a in high-power mode, as shown in Figure 13, but it is not limited to this. The high-power battery module 148 may supply incremental power based on the difference between the maximum required output and the baseline power to the first propulsion assembly 108a. In other words, the incremental power may be power that is less than or equal to the difference between the required output and the baseline power.

[0149] Processor 156 may determine the voltage applied to the first propulsion assembly 108a (the voltage of Node 1) in high-power mode. FDC 166 may control the input voltage of the fuel cell module 146 to the first propulsion assembly 108a based on the determined voltage. The input voltage control can be performed similarly during both take-off and landing flights. Additionally, the output voltage of the high-power battery module 148 can be controlled based on the determined voltage.

[0150] In some examples, upon detecting the high-power mode, processor 156 may supply the power output respectively from the high-power battery module 148 and the high-energy battery module 150 to the second propulsion assembly 110a. Processor 156 may determine the power output of the high-power battery module 148 and the high-energy battery module 150.

[0151] The power of the high-energy battery module 150 is determined based on the baseline power of the second propulsion assembly 110a, allowing the high-energy battery module 150 to support the power required by the second inverter 160 in high-power mode. If the first and second propulsion assemblies 108a and 110a have different performance / specifications or require different control based on settings, the baseline power of the second propulsion assembly 110a may be set independently of the baseline power of the first propulsion assembly 108a. The baseline power of the second propulsion assembly 110a may, for example, be set based on the power of the second propulsion assembly 110a in low-power mode, or in another example, it may be set independently, regardless of the low-power mode. If the first and second propulsion assemblies 108a and 110a have the same performance / specifications or require the same control in the same settings, the baseline power of the second propulsion assembly 110a may be set based on the baseline power of the first propulsion assembly 108a.

[0152] Processor 156 may determine the power of the high-energy battery module 150 to be below the baseline power of the second propulsion assembly 110a based on the flight state and the state of each battery module. Here, the baseline power of the second propulsion assembly 110a may either be independent or dependent on the baseline power of the first propulsion assembly 108a. For convenience of explanation, the present disclosure illustrates the high-energy battery module 150 as supplying the baseline power of the first propulsion assembly 108a to the second propulsion assembly 110a, as shown in Figure 14.

[0153] The power supplied by the high-power battery module 148 to the second propulsion assembly 110a may be determined based on the difference between the required output of the second propulsion assembly 110a in high-power mode and the baseline power of the first propulsion assembly 108a. Here, the required output of the first propulsion assembly 108a may be set as the maximum required output of the first propulsion assembly 108a in high-power mode, as described above, but it is not limited to this. The high-power battery module 148 may supply incremental power based on the difference between the maximum required output and the baseline power to the second propulsion assembly 110a. In other words, the incremental power may be power that is less than or equal to the difference between the required output and the baseline power.

[0154] Processor 156 may determine the voltage applied to the second propulsion assembly 110a (the voltage of Node 2) in high-power mode based on the charging state of the high-energy battery module 150. The high-power battery module 148 may control the voltage and current applied to the second propulsion assembly 110a using BHDC 170 based on the determined voltage. Additionally, the output voltage of the high-energy battery module 150 can be controlled based on the determined voltage. The control described above can be performed similarly during both take-off and landing flights.

[0155] FIG. 16 is a diagram showing power supply to the propulsion assembly in low-power mode with three multiple power sources operating in a normal state. Here, the cruise mode is exemplified as the flight mode applied during the cruise phase shown in Figures 13 and 14. Cruise mode is a type of low-power mode, and Figure 16 shows the power control of the battery in low-power mode.

[0156] Processor 156 may determine whether the system has entered cruise mode (or low-power mode) based on at least one of the flight control command and sensor data. In the present disclosure, the flight control command is a control request signal from the pilot or processor 156, for example, a request signal related to cruise flight or horizontal flight. The sensor data may be output from the sensor unit 132 which detects state in at least one of the motor, inverter, flight attitude change, and altitude change.

[0157] Processor 156 may tilt the first propulsion assembly 108a into a cruise arrangement and deactivate the second propulsion assembly 110a based on the control request signal. When the first propulsion assembly 108a operates in the cruise arrangement, processor 156, including FCC and PCU, may generate a current reduction command based on a specific output command and transmit it to the fuel cell converter FDC 166 and the bidirectional converter 170. The fuel cell converter 166 and bidirectional converter 170 may control the current applied to the first and second propulsion assemblies 108a and 110a based on the current reduction command. Processor 156 may determine that the system has entered low-power mode in response to the current reduction command, applied current, and tilt control. If the required outputs of the first propulsion assembly 108a during the transition phase when cruising begins at the target altitude are below the baseline power, processor 156 may maintain the low-power mode.

[0158] During the cruise flight before landing, when the air mobility apparatus 100 is flying near the destination, processor 156 may reduce the horizontal speed based on the control request signal. Additionally, processor 156, including FCC, may generate a current reduction command based on the control request signal. Furthermore, air mobility apparatus 100 can operate in inertial flight, consuming less energy than during the previous cruise flight by utilizing the speed of cruise flight. Accordingly, the fuel cell module 146 may supply less power than the baseline power to the first propulsion assembly 108a, and the remaining power can be transferred to other battery modules 148 and 150. Processor 156 may maintain the system in low-power mode in response to the maximum reduction command, applied current, tilt control, and inertial flight operation.

[0159] Apart from the above-described examples, the details regarding entering high-power mode are described in Figure 6 and will be omitted here.

[0160] Upon detecting cruise mode (i.e., low-power mode), processor 156 may supply the power output from the fuel cell module 146 to the first propulsion assembly 108a. Processor 156 may determine the power output of the fuel cell module 146.

[0161] The power output of the fuel cell module 146 may be determined based on the baseline power of the first propulsion assembly 108a in low-power mode, as illustrated in Figure 15. The high-power battery module 148 may not supply power to the first propulsion assembly 108a and may operate in a dormant state. The detailed explanation of the baseline power is provided in Figure 15 and will be omitted here.

[0162] Since the second propulsion assembly 110a is mainly deactivated in cruise mode, processor 156 may stop the power supply to the second propulsion assembly 110a by the high-power battery module 148 and the high-energy battery module 150. If the flight status in cruise mode consumes energy above a set threshold greater than the baseline power, processor 156 may activate the second propulsion assembly 110a using the fuel cell module 146.

[0163] Processor 156 may check the charging status of the high-power battery module 148 and the high-energy battery module 150 in response to the low-power mode. Processor 156 may determine whether to charge the high-power battery module 148 and the high-energy battery module 150 based on the checked charging status. The charging status may include at least one of the current charge level and the required charge level. The required charge level may be the charging state required based on the flight environment, for example. The flight environment may include the weather at the destination, an unexpected delay in the landing schedule, changes in the flight route, and weather along the flight route, but is not limited to these factors.

[0164] In response to the charging decision, processor 156 may cause the fuel cell module 146 to charge at least one of the high-power battery module 148 and the high-energy battery module 150. While transitioning from the low-power mode to the high-power mode, the fuel cell module 146 may charge at least one of the high-power battery module 148 and the high-energy battery module 150. In some examples, The fuel cell module 146 may charge the high-power battery module 148 and the high-energy battery module 150 during inertial flight (as illustrated in the inertial flight phase in Figure 13). The excess power can be used for charging by subtracting the actual power consumed by the propulsion assembly during inertial flight from the baseline power. Processor 156 may start charging the battery modules by the fuel cell module 146 before the inertial flight phase according to the flight status. The target charge level for each battery module 148 and 150 may be, for example, a pre-designated charge level, or alternatively, it may be dynamically determined based on the power required for landing hover and the flight conditions.

[0165] The fuel cell module 146 may charge at least one of the high-power battery module 148 and the high-energy battery module 150 and supply power to the first propulsion assembly 108a through FDC 166 and BHDC 170. Accordingly, power source redundancy can be realized, and powertrain weight reduction can be achieved by carrying only the minimum battery modules for take-off, without considering the power required for landing.

[0166] FIG. 17 is a diagram showing an example of redundancy control responding to failure in three multiple power sources. The present implementation illustrates redundancy control in response to a failure state of the fuel cell module 146. Hereinafter, the three power sources may correspond to the first to third battery modules, as previously described.

[0167] The processor 156 may detect the status of all power sources 146, 148, and 150 during flight through the sensor unit 132. At least one of the voltage / current sensor 132a, temperature sensor 132b, or resistance sensor 132d may be installed within the power source, at the output terminal of the power source, inside or at the output terminal of the fuel cell converter 166, or within the bidirectional converter 170.

[0168] In the present disclosure, the status may be a normal or failure state. If the voltage or current of the fuel cell module 146 or each battery module 148, 150 exceeds a threshold value, or if the resistance of these modules falls outside the normal range, the processor 156 may determine that the corresponding module is in a failure state. Alternatively, if a predetermined current is not detected at the fuel cell converter 166, which is connected to and operates the fuel cell module 146 based on the flight mode, the processor 156 may determine that the fuel cell module 146 is in a failure state. If a predetermined current is not detected at the bidirectional converter 170, which is connected to the battery modules 148, 150 based on the flight mode, the processor 156 may determine that the corresponding module is in a failure state.

[0169] Additionally, the status may include the charge status, output voltage, output current, generation status, remaining fuel of the fuel cell module 146, the operating status of the fuel cell converter 166, and the operating status of the bidirectional converter 170.

[0170] With respect to FIG. 17, in the event of a failure in the fuel cell module 146, the processor 156 may deactivate the fuel cell module 146. Additionally, the processor 156 may identify power sources that operate in a normal state. The power sources that operate in a normal state are exemplified as the high-power battery module 148 and the high-energy battery module 150.

[0171] As described in FIGS. 15 and 16, the fuel cell module 146 may be a power source that operates in both the high-power mode and the low-power mode. The redundancy control in FIG. 17 is common to both the high-power mode and low-power mode, so it may be described without distinguishing between flight modes. However, if the redundancy control is related to a specific flight mode, the flight mode will be specified.

[0172] If the fuel cell module 146 is in a failure state, the processor 156 may be configured to supply the respective power from the high-power battery module 148 and the high-energy battery module 150 to the first propulsion assembly 108a, based on at least one of required information and status information.

[0173] Furthermore, the processor 156 may deactivate the power supply from the high-power battery module 148 to the second propulsion assembly 110a. The processor 156 may be configured to supply power from the high-energy battery module 150 to the second propulsion assembly 110a, based on the required output of the first and second propulsion assemblies 108a, 110a according to the flight mode, and the status information. Since the high-energy battery module 150 has a lower output power but higher energy capacity compared to the high-power battery module 148, the high-energy battery module 150 may supply power to the first and second propulsion assemblies 108a, 110a to ensure safe flight for as long as possible.

[0174] In addition, charging the battery modules 148, 150 by the fuel cell module 146 during operation in the low-power mode may be suspended.

[0175] The required information may be the required output of the first and second propulsion assemblies 108a, 110a, based on the flight mode.

[0176] The status information may include the status of the normal high-power battery module 148, the status of the normal high-energy battery module 150, and the flight status. The status of the high-power battery module 148 and the high-energy battery module 150 may include, for example, the charge status, sensor data for the batteries (such as temperature, current, voltage), anticipated battery demand based on the flight path and loading conditions, and the charging history. However, the status information is not limited to these examples and may include a variety of data related to battery conditions. In the implementations shown in FIGS. 18 and 19, the status of the fuel cell module 146 may include, for example, the remaining fuel status of the fuel cell module 146, generation status, anticipated battery demand, and the charging history for the battery modules 148, 150, without being limited to these examples.

[0177] The flight status may include the flight plan, surrounding weather conditions, and flight operating conditions. The flight plan may include, for example, the remaining distance to the destination, the remaining flight time, and the difficulty of flight control along the flight path. The destination may be the final destination or an intended waypoint. The flight operating conditions may include, for example, the load weight, the arrangement of cargo within the aircraft, flight time, the total flight path, and the required flight time, without being limited to these examples.

[0178] In one example, in the case of a high-power mode during landing, the processor 156 may determine whether the state of the normal battery modules 148 and 150 can support the required output of the first and second propulsion assemblies 108a and 110a. If support is possible, the processor 156 may configure the power supplied from each battery module 148 and 150 to the first and second propulsion assemblies 108a and 110a based on the required output of the first and second propulsion assemblies 108a and 110a in the high-power mode. For example, the high-energy battery module 150 may output baseline power to the first propulsion assembly 108a, and the high-power battery module 148 may output incremental power to the first propulsion assembly 108a. The baseline power for the first propulsion assembly 108a may correspond to the power of the normal state of the fuel cell module 146. The high-energy battery module 150 may provide the required output of the second propulsion assembly 110a or a value close to it to the second propulsion assembly 110a. Without being limited to these examples, the processor 156 may adaptively determine the power from each battery module 148 and 150 to be supplied to the propulsion assemblies 108a and 110a based on the status of each normal battery module.

[0179] In another example, in the case of a low-power mode during cruise flight, the processor 156 may determine whether the remaining distance to the destination and the status of the normal battery modules 148 and 150 allow flight to the destination. The processor 156 may analyze whether the required output of each propulsion assembly 108a and 110a during cruise flight and landing can be supported by the status of the normal battery modules. If support is possible, the processor 156 may configure the power supplied from each battery module 148 and 150 to the propulsion assemblies 108a and 110a based on the required output of the first and second propulsion assemblies 108a and 110a during the low-power mode of cruise flight before landing. In this case, the combined power of the high-power battery module 148 and the high-energy battery module 150 may provide the baseline power to the first propulsion assembly 108a. The high-energy battery module 150 may provide the baseline power to the second propulsion assembly 110a. The baseline power for the second propulsion assembly 110a may correspond to the power of the normal state of the fuel cell module 146. In the high-power mode during landing, the processor 156 may determine the power of each battery module 148 and 150 according to the description provided for the high-power mode during landing and supply it to each propulsion assembly 108a and 110a. The power of each battery module is not limited to the examples described above and may be adaptively determined based on the required information and status information.

[0180] In another example, in the case of a low-power mode during cruise flight or a high-power mode during takeoff, the processor 156 may determine whether flight to the destination is possible based on the remaining distance and the status of the normal battery modules 148 and 150. The processor 156 may analyze whether the required output of each propulsion assembly 108a and 110a for flight to the destination can be supported by the status of the normal battery modules 148 and 150. In addition, the processor 156 may confirm whether the air mobility apparatus 100 is flying at an initial point along the flight path or whether an emergency landing request has been issued by the control server (control tower server). The initial point along the flight path may correspond to a location within a certain percentage of the total flight path.

[0181] The processor 156 may determine the flight mode to be an emergency landing mode based on at least one of the inability to fly to the destination, flight at the initial point, or the emergency landing request. The emergency landing mode may request the aircraft to land at the departure point, a waypoint already passed, or a nearby landing field. The processor 156 may supply power from the normal battery modules 148 and 150 to the first and second propulsion assemblies 108a and 110a in accordance with the emergency landing mode. Since the emergency landing mode involves both cruise flight and landing, the processor 156 may determine the power supplied from each battery module 148 and 150 to the propulsion assemblies 108a and 110a, similar to the low-power mode example for cruise flight. The power may be adaptively adjusted based on the battery status and flight conditions of the battery modules 148 and 150 and may be less than the power mentioned in the low-power mode example. Accordingly, the air mobility apparatus 100 may land safely, powered by a stable output for a certain period of time.

[0182] FIG. 18 is a diagram showing the operation of a bidirectional converter.

[0183] If the fuel cell module 146 is in a failure state, the voltage at node 1 may be determined by the charge status of the high-power battery module 148. The voltage at node 2 may be determined by the charge status of the high-energy battery module 150. The high-energy battery module 150 may supply power to the first propulsion assembly 108a through voltage and current control in the BHDC 170. At this time, the input-output relationship of the BHDC 170 may be expressed as V_node1 = V_node2(1-D') / 1. D' may refer to the turn-on duty ratio of the high-side switch 178 of the BHDC.

[0184] The redundancy control in FIG. 17 illustrates that the high-power battery module 148 supplies power only to the first propulsion assembly 108a. In another example, if the battery capacity or charge status of the high-power battery module 148 can meet the required output of the first and second propulsion assemblies 108a and 110a according to the flight mode, the high-power battery module 148 may supply additional power to the second propulsion assembly 110a based on the required information and status information.

[0185] FIG. 19 is a diagram showing another example of redundancy control responding to failure in three multiple power sources. The present implementation illustrates redundancy control in response to a failure in the high-power battery module 148.

[0186] When the processor 156 detects a failure in the high-power battery module 148, the processor 156 may use a protection element 174 to cut off the output of the high-power battery module 148. Additionally, the processor 156 may identify a power source that is operating in a normal state. The power sources operating normally are illustrated as the fuel cell module 146 and the high-energy battery module 150.

[0187] As explained in FIGS. 15 and 16, the high-power battery module 148 may operate in a high-power mode. Unless otherwise specified, the flight mode in which the redundancy operates may not be described.

[0188] The processor 156 may respond to the failure in the high-power battery module 148 by supplying power from the fuel cell module 146 and the high-energy battery module 150 to the first propulsion assembly 108a based on at least one of the required information and status information. The required information and status information described in FIG. 17 can be interpreted in a manner consistent with the present information. For instance, the status of a normal battery module (or a normal power sources) in the status information may be understood as the status of the normal fuel cell module 146 and the normal high-energy battery module 150. The processor 156 may be configured to supply power greater than or equal to the power of the normal state from the fuel cell module 146 to the first propulsion assembly 108a. The power of the normal state may correspond to the baseline power of the first propulsion assembly 108a.

[0189] Moreover, the processor 156 may control the fuel cell module 146 to supply power solely to the first propulsion assembly 108a. The processor 156 may determine the power supplied from the high-energy battery module 150 to the second propulsion assembly 110a based on the required output of the first and second propulsion assemblies 108a and 110a, the status of the power of the normal power sources 146 and 150, and the flight status, ensuring that the power is supplied to the second propulsion assembly 110a. As the high-energy battery module 150 has higher output power compared to the fuel cell module 146, the high-energy battery module 150may transfer power to both the first and second propulsion assemblies 108a and 110a to ensure safe flight.

[0190] When the high-power battery module 148 is in the failure state during a low-power mode, the fuel cell module 146 may stop charging the high-power battery module 148. Even if the air mobility apparatus 100 is flying in a low-power mode, the processor 156 may determine whether to execute the redundancy control described in FIG. 19 based on at least one of the flight mode, required information, and status information.

[0191] For instance, in the case of a high-power mode during landing, the processor 156 may determine whether the status of the power of the normal power sources 146 and 150 can support the required output of the first and second propulsion assemblies 108a and 110a. If support is possible, the processor 156 may set the power supplied from each power source 146 and 150 to the first and second propulsion assemblies 108a and 110a based on the required output of the first and second propulsion assemblies 108a and 110a in the high-power mode. For example, the fuel cell module 146 may output baseline power to the first propulsion assembly 108a, and the high-energy battery module 150 may output incremental power to the first propulsion assembly 108a. The high-energy battery module 150 may provide the required output of the second propulsion assembly 110a or a value close to it to the second propulsion assembly 110a. Without being limited to these examples, the processor 156 may adaptively determine the power from each power source based on the status of each power of the normal source 146 and 150 to supply power to the propulsion assemblies 108a and 110a.

[0192] In another example, during a low-power mode in cruise flight, the processor 156 may determine whether flight to the destination is possible based on the remaining distance and the status of the power of the normal power sources 146 and 150. The determination process may be similar to that explained in FIG. 17. If flight to the destination is possible, the processor 156 may supply power from the fuel cell module 146 to the first and second propulsion assemblies 108a and 110a to enable cruise flight before landing, as described in FIG. 16. In the case of a high-power mode during landing, the processor 156 may determine the power of each power source 146 and 150 and supply it to each propulsion assembly 108a and 110a based on the aforementioned details. The power of each power source is not limited to the examples described and may be adaptively determined based on the required information and status information.

[0193] In another example, during a low-power mode in cruise flight or a high-power mode during takeoff, the processor 156 may determine whether the flight mode is an emergency landing mode similar to the explanation in FIG. 17. If the emergency landing mode is determined, the processor 156 may supply power from the power of the normal power sources 146 and 150 to the first and second propulsion assemblies 108a and 110a in accordance with the emergency landing mode. The power control in the emergency landing mode may be similar to that described in FIG. 17.

[0194] In some examples, if the high-power battery module 148 is in a failure state, the voltage of the high-energy battery module 150 may be converted into power of the first propulsion assembly 108a through the BHDC 170 and transmitted to the first propulsion assembly 108a. For power transmission, the voltage required by the first propulsion assembly 108a may be generated through the BHDC 170. By controlling the input voltage, the FDC may output the voltage at node 1, allowing the power generated by the fuel cell module 146 to be supplied to the first propulsion assembly 108a. To generate the voltage of node 1, namely, the voltage required by the first propulsion assembly 108a, power conversion may be performed via the BHDC 170. In this case, the input-output relationship of the BHDC can be expressed as V_node1 = (1-D') / V_node2, where D' is the turn-on duty ratio of the high-side switch 178 of the BHDC.

[0195] The redundancy control described in FIG. 19 illustrates that the fuel cell module 146 supplies power only to the first propulsion assembly 108a. In another example, if the output power of the fuel cell module 146 can meet the required output of the first and second propulsion assemblies 108a and 110a according to the flight mode, the fuel cell module 146 may supply additionally power to the second propulsion assembly 110a based on the required information and status information.

[0196] FIG. 20 is a diagram showing yet another example of redundancy control responding to failure in three multiple power sources. The present implementation illustrates redundancy control in response to a failure in the high-energy battery module 150.

[0197] When the processor 156 detects a failure state in the high-energy battery module 150, the processor 156 may use a safety element 176 to cut off the output of the high-energy battery module 150. Additionally, the processor 156 may identify a power source that is operating in the normal state The power sources operating normally are exemplified as the fuel cell module 146 and the high-power battery module 148.

[0198] As explained in FIGS. 15 and 16, the high-energy battery module 150 may operate in a high-power mode. Unless otherwise specified, the flight mode in which the redundancy operates may not be described.

[0199] The processor 156 may be configured to supply power from each of the fuel cell module 146 and the high-power battery module 148 to the first and second propulsion assemblies 108a and 110a in response to a failure of the high-power battery module 148, based on at least one of the flight mode, required information, and status information. The required information and status information described in FIG. 17 may be interpreted in accordance with the information outlined above. For example, the status of a normal battery module (or the normal power source) in the status information may be understood as the status of the normal fuel cell module 146 and the normal high-power battery module 148. The processor 156 may be configured to supply power greater than or equal to the power of the normal state from the fuel cell module 146 to the first propulsion assembly 108a. The power of the normal state may correspond to the baseline power of the first propulsion assembly 108a.

[0200] Since the high-power battery module 148 has higher output power but lower energy capacity compared to the fuel cell module 146, the fuel cell module 146 may deliver power to the first and second propulsion assemblies 108a and 110a to ensure safe flight for as long as possible.

[0201] On the other hand, if the high-energy battery module 150 fails during low-power mode, the fuel cell module 146 may stop charging the high-energy battery module 150. Additionally, even if the air mobility apparatus 100 is flying in low-power mode, the processor 156 may determine whether to execute the redundancy control described in FIG. 20, based on at least one of the flight mode, required information, and status information.

[0202] For instance, in the case of high-power mode during landing, the processor 156 may determine, in a manner similar to that of FIG. 19, whether the status of the power of the normal power sources 146 and 148 can support the required output of the first and second propulsion assemblies 108a and 110a. If support is possible, the processor 156 may set the power supplied from each power source 146 and 148 to the first and second propulsion assemblies 108a and 110a based on the required output of the first and second propulsion assemblies 108a and 110a in high-power mode. For example, the fuel cell module 146 may output baseline power to the first and second propulsion assemblies 108a and 110a, and the high-power battery module 148 may output incremental power to the first and second propulsion assemblies 108a and 110a. Without being limited to these examples, the processor 156 may adaptively determine the power supplied from each battery module 148 and 150 to the propulsion assemblies 108a and 110a based on the status of the power of the normal power sources.

[0203] In another example, in the case of low-power mode during cruise flight, the processor 156 may determine whether flight to the destination is possible based on the remaining distance and the status of the power of the normal power sources 146 and 150. The decision process may be similar to the one described in FIG. 17. If flight to the destination is possible, the processor 156 may perform cruise flight before landing, using power control as described in FIG. 16. In the case of high-power mode during landing, the processor 156 may determine the power of each power source 146 and 150 and supply it to each propulsion assembly 108a and 110a based on the aforementioned matters. The power of each power source is not limited to the examples described and may be adaptively determined based on the required information and status information.

[0204] In another example, in the case of low-power mode during cruise flight or high-power mode during takeoff, the processor 156 may determine whether the flight mode is emergency landing mode, similar to the explanation in FIG. 17. If emergency landing mode is determined, the processor 156 may supply power from the power of the normal power sources 146 and 148 to the first and second propulsion assemblies 108a and 110a in accordance with the emergency landing mode. Power control in the emergency landing mode may be similar to that described in FIG. 17.

[0205] In some examples, if the high-energy battery module 150 is in a failure state, the FDC 166 may control the input voltage to set the voltage of the high-power battery module 148 as the output voltage and apply it to the first propulsion assembly 108a. The BHDC 170 may convert the output voltage to the voltage required by the second propulsion assembly 110a and transfer the power from the fuel cell module 146 and high-power battery module 148 to the second propulsion assembly 110a. To generate the voltage at node 2, or the voltage required by the second propulsion assembly 110a, power conversion may be performed through the BHDC. In this case, as referenced in FIG. 18, the input-output relationship of the BHDC may be expressed as V_node2 = V_node1 / (1-D), where D is the turn-on duty ratio of the low-side switch 180 of the BHDC.

[0206] The redundancy control described in FIG. 20 illustrates that the fuel cell module 146 and high-power battery module 148 supply power to all propulsion assemblies 108a and 110a. In another example, the fuel cell module 146 and high-power battery module 148 may be controlled to supply power to the first and second propulsion assemblies 108a and 110a in various power distribution configurations based on the flight mode, required information, and status information.

[0207] FIG. 21 is a diagram showing power supply to the propulsion assembly in high-power mode with two multiple power sources operating in a normal state. The implementations of Figures 19 and 20 differ from Figure 10 in terms of the configuration and power control of multiple power sources. The first propulsion assembly 108a and the second propulsion assembly 110a include a tiltable rotor and a fixed rotor, respectively, and their details are substantially the same as those in Figures 10 to 12. The following description will focus on the differences from the implementation of Figure 10.

[0208] Figs. 21 and 22 show that the power source unit 142 includes two batteries. The power source unit 142, for example, a fuel cell module 246 and a high-power battery module 250. The fuel cell module 246 and the high-power battery module 250a, 250b correspond to the first and second battery modules, respectively. For the sake of convenience, the implementation where the power source unit 142 includes the aforementioned battery combination will be described. In another example, the power source unit 142 may include the fuel cell module 246 and a high-energy battery module 150.

[0209] Although Figs. 21 and 22 illustrate the first propulsion assembly 108a and the second propulsion assembly 110a as being single units, the implementations of Figs. 21 and 22 can be commonly applied to other first and second propulsion assemblies (108b-108d, 110b-110d). The high-power battery module 250 is not limited to one unit and may include multiple battery modules connected to other first and second propulsion assemblies (108b-108d, 110b-110d). Fig. 21 illustrates that the high-power battery module includes the first high-power battery module 250a and the second high-power battery module 250b.

[0210] The fuel cell module 246 and multiple high-power battery modules 250a, 250b can be connected in parallel to the common node A. The fuel cell module 246 can be selectively connected to each high-power battery module 250a, 250b to supply charging power to them. This selective connection can be implemented by the switching elements of each high-power battery module 250a, 250b. The fuel cell module 246 can be electrically connected to the first propulsion assembly 108a and the second propulsion assembly 110a.

[0211] Each of the first and second high-power battery modules 250a, 250b may have homogeneous characteristics. Depending on the flight mode, each of the first and the second high-power battery module 250a, 250b can be selectively connected to the first and second propulsion assemblies 108a and 110a. This selective connection can be carried out by the switching elements of the high-power battery module 250a, 250b. Similar to Figure 10, the fuel cell module 246 may include a control device implemented as a power management controller (PMC). The fuel cell converter 266 and reverse current prevention device 268 may be arranged between the fuel cell module 246 and node A. These elements function in the same manner as described in Figures 10 and 11.

[0212] As shown in Fig. 21, the first and second bidirectional converters 270a and 270b may be placed between each high-power battery module 250 and node A. The bidirectional converters 270a and 270b can be, for example, BHDCs, and their detailed operation is described in Figures 10 and 11.

[0213] As shown in Fig. 21, when hover mode (i.e., high-power mode) is detected, processor 156 may supply the power output from the fuel cell module 246 to the first and second propulsion assemblies 108a and 110a. Additionally, each high-power battery module 250 can supply power to the corresponding propulsion assembly, i.e., the first and second propulsion assemblies 108a and 110a. The high-power mode during takeoff or landing flight can be detected in the manner described in Figure 15. Processor 156 can determine the power to be output by the fuel cell module 246 and each of the multiple high-power battery modules 250.

[0214] The power from the fuel cell module 246 may be determined based on the baseline power of at least one of the first and second propulsion assemblies 108a and 110a. The baseline power may be set based on the power applied to at least one of the first and second propulsion assemblies 108a and 110a in low-power mode. Here, the baseline power of the second propulsion assembly 110a can be independent of or dependent on the baseline power of the first propulsion assembly 108a. Processor 156 may determine the power of the fuel cell module 246 to be below the baseline power based on the flight status and the state of each battery module. In the present disclosure, the fuel cell module 246 supplies baseline power to the first and second propulsion assemblies 108a and 110a, as illustrated in Figures 13 and 14. As shown in Figures 13 and 14, the baseline power of the second propulsion assembly 110a may be the same as that of the first propulsion assembly 108a, allowing the fuel cell module 246 to supply the baseline power of the first propulsion assembly 108a to both the first and second propulsion assemblies 108a and 110a. Additionally, the power supplied by the fuel cell module 246 to each of the propulsion assemblies 108a and 110a can be the same in both high-power and low-power modes.

[0215] The power supplied by the first high-power battery module 250a to the first propulsion assembly 108a can be determined, similarly to Figure 15, based on the difference between the required output of the first propulsion assembly 108a in high-power mode and its baseline power. The required output of the first propulsion assembly 108a can be set as the maximum required output in high-power mode, as in Figure 13, but is not limited to this. The first high-power battery module 250a can supply the incremental power based on the difference between the maximum required output and the baseline power of the first propulsion assembly 108a. The power supplied by the second high-power battery module 250b to the second propulsion assembly 110a may be determined similarly, based on the required output and incremental power of the second propulsion assembly, as illustrated in Figure 14.

[0216] In high-power mode, processor 156 can determine the voltage applied to each propulsion assembly 108a and 110a and control the input voltage of the fuel cell module 246 to each propulsion assembly 108a and 110a based on the determined voltage. The input voltage control can be performed by the FDC 266. Additionally, the current supplied to each of the first propulsion assembly 108a and the second propulsion assembly 110a can be controlled based on the voltage applied to the first propulsion assembly 108a. Here, the currents supplied to the first propulsion assembly 108a and the second propulsion assembly 110a by the fuel cell module 246 can be set to be the same. Each current may be the current flowing from the fuel cell module 246 to node A on the first propulsion assembly side and from the fuel cell module 246 to node A on the second propulsion assembly side.

[0217] The first BHDC 270a functions as a primary device and can control the output voltage at node A. The voltage at node A may be output by the first BHDC 270a to be greater than the back electromotive force of the first propulsion assembly 108a at maximum output. Consequently, the current flowing from the FDC 266 to the first propulsion assembly 108a (i.e., to node A on the first propulsion assembly side) may be determined in a dependent manner. The second BHDC 270b functions as a secondary device and controls the output current.

[0218] Specifically, to generate the voltage required by the first propulsion assembly 108a during takeoff or landing, the first BHDC 270a generates and controls the output voltage at node A. The total current output by the FDC 266 may be determined by dividing the fuel cell output by the voltage at node A such that the FDC 266 controls the input voltage based on the required output of the fuel cell module 246. The current flowing from the FDC 266 to the first propulsion assembly 108a may be determined by the output current of the second propulsion assembly 110a. The second BHDC 270b can control the current so that the current flowing from the FDC 266 to the first propulsion assembly 108a and the current flowing from the FDC 266 to the second propulsion assembly 110a are the same. If the required current (or output current) of the second propulsion assembly 110a differs from that of the first propulsion assembly 108a, the second BHDC 270b can control the current based on the required current of the second propulsion assembly 110a. Here, the required current of the second propulsion assembly 110a may be determined based on the current flowing from the FDC 266 to the second propulsion assembly 110a and the output current of the second BHDC 270b. For example, the required current of the second propulsion assembly 110a may be the sum of the aforementioned currents.

[0219] FIG. 22 is a diagram showing power supply to the propulsion assembly in low-power mode with two multiple power sources operating in a normal state.

[0220] Upon detecting cruise mode (i.e., low-power mode), processor 156 may supply the power output from the fuel cell module 246 to the first propulsion assembly 108a. Processor 156 can determine the power to be output by the fuel cell module 246. The low-power mode according to the cruise flight can be detected in the manner described in Figure 16.

[0221] The power of the fuel cell module 246 may be determined based on the baseline power of the first propulsion assembly 108a in low-power mode, as illustrated in Figure 15. The first high-power battery module 250a may not supply power to the first propulsion assembly 108a and may operate in a dormant state. The detailed description of the baseline power is omitted, as it has already been explained.

[0222] Since the second propulsion assembly 110a is primarily deactivated in cruise mode, processor 156 may stop the power supply to the second propulsion assembly 110a from the second high-power battery module 250b. In cruise mode, the fuel cell module 246 may supply power to the second propulsion assembly 110a based on flight status, as controlled by processor 156. The detailed explanation of this is provided in Figure 16.

[0223] In response to low-power mode, the processor 156 may check the state of charge of each high-power battery module 250a, 250b. Based on the checked charge status, processor 156 can determine whether each high-power battery module 250a, 250b needs to be charged. The state of charge may include at least one of the current charge level and the required charge level. The required charge level may, for instance, be the state of charge based on the flight environment. The details of the above is described in Fig. 16.

[0224] Processor 156 can instruct the fuel cell module 246 to charge at least one of the multiple high-power battery modules 250a, 250b in response to the charging decision. While transitioning from the low-power mode to the high-power mode, the fuel cell module 246 may charge at least one of the high-power battery modules 250a, 250b. In some examples, The fuel cell module 246 can charge the high-power battery module 250 during inertia flight. The details of the charge amount, charge timing, and charge control are substantially the same as those described in Figure 16.

[0225] The first BHDC 270a determines the voltage of node A, and the FDC 266 performs input voltage control so that the voltage of node A serves as the output voltage. Since the FDC 266 supplies power generated by the fuel cell module 246 to the first and second propulsion assemblies 108a, 110a and the output of the fuel cell module 246 is higher than the output used by the first and second propulsion assemblies 108a and 110a, the fuel cell module 246 can charge at least one of the high-power battery modules 250a, 250b. The second BHDC 270b performs negative current control based on the voltage of node A and carries out the operation of charging the high-power battery module 250. Here, the current flowing from the FDC 266 to node A on the second propulsion assembly side can be defined as the sum of the current flowing from node A to the second propulsion assembly 110a and the current flowing from node A to the second BHDC 270b.

[0226] FIG. 23 is a diagram showing an example of redundancy control responding to failure in two multiple power sources. The present implementation illustrates redundancy control corresponding to a failure state of the fuel cell module 246. Here, the two multiple power sources may correspond to the first and second battery modules as described above.

[0227] When the processor 156 detects a failure state in the fuel cell module 246, the processor 156 may cut off the output of the fuel cell module 246. The power sources operating in a normal state are exemplified as the plurality of high-power battery modules 250a and 250b.

[0228] As described in FIGS. 21 and 22, the fuel cell module 246 may drive in both high-power mode and low-power mode. Since the redundancy control of FIG. 23 is common to both high-power mode and low-power mode, it may be described without distinguishing between flight modes. However, if redundancy control is related to a specific flight mode, the flight mode may be explicitly stated.

[0229] The processor 156 may be configured to supply power from each of the plurality of high-power battery modules 250a and 250b to the first and second propulsion assemblies 108a and 110a based on at least one of the flight mode, required information, and status information in response to the failure state of the fuel cell module 246. The required information and status information described in FIGS. 17 to 20 may be interpreted in accordance with the information outlined above. In some examples, the fuel cell module 246 may stop charging the plurality of high-power battery modules 250a and 250b.

[0230] For example, in the case of high-power mode during landing, the processor 156 may determine the power to be supplied from the first and second high-power battery modules 250a and 250b to the first and second propulsion assemblies 108a and 110a, based on at least one of the required output of each propulsion assembly 108a and 110a in high-power mode, the state of the battery modules 250a and 250b, and the flight status. The determined powers may be equal to or less than the maximum required output and may be the same for both. This is limited to this example, the processor 156 may set different power levels for each propulsion assembly 108a and 110a or adaptively determine the power supplied from the plurality of high-power battery modules 250a and 250b, based on the state of the normal battery modules.

[0231] In another example, in the case of low-power mode during cruise flight, the processor 156 may determine whether flight to the destination is possible based on the remaining distance, the state of the normal high-power battery modules 250a and 250b, and the power of the normal state of the fuel cell module 246. The decision process may be similar to the method described in FIG. 17. If flight to the destination is possible, the processor 156 may set the power supplied from each high-power battery module 250a and 250b during cruise flight to be the same as the power of the normal state of the fuel cell module 246 in low-power mode. In the case of high-power mode during landing, the processor 156 may determine the power for each of the plurality of high-power battery modules 250a and 250b and supply that power to each propulsion assembly 108a and 110a in accordance with the aforementioned matters. The power of each high-power battery module is not limited to the examples described and may be adaptively determined based on the required information and status information.

[0232] The processor 156 may determine whether the flight mode is an emergency landing mode, similar to the explanation in FIG. 17. If emergency landing mode is determined, the processor 156 may supply the power from the normal high-power battery modules 250a and 250b to the first and second propulsion assemblies 108a and 110a in accordance with the emergency landing mode. Power control in emergency landing mode may be similar to that described in FIG. 17.

[0233] In some examples, if the fuel cell module 246 is in a failure state, the first BHDC 270a, functioning as the primary device, may perform voltage control, and the second BHDC 270b, functioning as the secondary device, may perform current control, similar to the normal operation of the plurality of high-power battery modules 250a and 250b. Thus, the power from the high-power battery modules 250a and 250b may be delivered to the first and second propulsion assemblies 108a and 110a.

[0234] The redundancy control in FIG. 23 exemplifies the first and second high-power battery modules 250a and 250b each supplying power only to the first and second propulsion assemblies 108a and 110a, respectively. In another example, if the output power of each high-power battery module 250a and 250b can correspond to the required output of another propulsion assembly, the high-power battery modules 250a and 250b may supply additional power to another propulsion assembly 108a and 110a, based on the required information and status information.

[0235] FIG. 24 is a diagram showing another example of redundancy control responding to failure in two multiple power sources. The present implementation illustrates redundancy control corresponding to a failure state in the first high-power battery module 250a, which is connected to the bidirectional converter 270a functioning as the primary device.

[0236] When the processor 156 detects a failure state in the first high-power battery module 250a, the processor 156 may use a protective device to cut off the output of the fuel cell module 246. The power sources operating in a normal state are exemplified as the fuel cell module 146 and the second high-power battery module 250b.

[0237] As described in FIGS. 21 and 22, the first high-power battery module 250a may operate in high-power mode. In the following, unless otherwise specified, the flight mode where redundancy operates may not be described.

[0238] The processor 156 may supply the power of the normal state from the fuel cell module 246 to the first and second propulsion assemblies 108a and 110a in response to the failure state of the first high-power battery module 250a. The power of the normal state may correspond to the baseline power of each propulsion assembly 108a and 110a.

[0239] For example, the processor 156 may determine the power of the second high-power battery module 250b, based on at least one of the incremental power of the first propulsion assembly 108a, the state of each power source 246 and 250b, and the flight status, and supply that power to the first propulsion assembly 108a. Additionally, the processor 156 may determine the power of the second high-power battery module 250b, based on at least one of the incremental power of the second propulsion assembly 110a, the state of each power source 246 and 250b, and the flight status, and supply that power to the second propulsion assembly 110a. If the first high-power battery module 250a fails in low-power mode, the fuel cell module 246 may stop charging the first high-power battery module 250a.

[0240] The flight mode may be at least one of high-power mode during landing, low-power mode during cruise flight, and emergency landing mode. The processor 156 may control power distribution for each flight mode in a manner similar to the examples described in FIGS. 19, 20, and 23.

[0241] In some examples, if the first high-power battery module 250a is in a failure state, the first high-power battery module 250a may not deliver power to the first BHDC 270a, functioning as the master, and therefore may not perform output voltage control. The processor 156 may switch the second BHDC 270b to function as the primary device. Accordingly, the function of the second BHDC 270b may be changed from output current control to output voltage control, enabling power control of the fuel cell module 146 and the second high-power battery module 250b.

[0242] The redundancy control in FIG. 24 illustrates that the fuel cell module 146 and the second high-power battery module 250b supply power to all propulsion assemblies 108a and 110a. In another example, at least one of the power sources may supply power to some of the propulsion assemblies 108a and 110a, based on the required information and status information.

[0243] The above explanation relates to redundancy control corresponding to a failure state in the first high-power battery module 250a. As another example, redundancy control corresponding to a failure state in the second high-power battery module 250b may be substantially the same as described above. However, since the second BHDC 270b connected to the second high-power battery module 250b functions as the secondary device, the first high-power battery module 250a and the fuel cell module 246 may deliver power to the propulsion assemblies 108a and 110a via the first BHDC 270a.

[0244] In this disclosure, the first and second propulsion assemblies are described as being different types of rotors. This disclosure can also be applied to air mobility apparatuses where all propulsion assemblies include the same type of rotor. Specifically, this disclosure can be applied to a tilting rotor type where all propulsion assemblies include tilting rotors, or a multicopter type where all propulsion assemblies include fixed rotors.

[0245] In the case of high-power mode, where all propulsion assemblies tilt or are fixed, the first battery (e.g., the fuel cell module) and the second battery (e.g., either the high-power battery module or the high-energy battery module) may each supply power to the propulsion assemblies. In low-power mode, the first battery may supply power to the propulsion assemblies. Additionally, the first battery may charge the second battery depending on the state of charge of the second battery.

[0246] Furthermore, in the type where all propulsion assemblies tilt or are fixed, in response to a failure state of the first battery, the power supplied from the second battery to each propulsion assembly may be determined based on at least one of the flight mode, required information, and status information. In response to a failure state of the second battery, power greater than or equal to the power of the normal state from the first battery may be supplied to each propulsion assembly. The matters described in the figures can be implemented in the aforementioned types of air mobility apparatuses as long as they do not contradict technically.

[0247] While the example methods of the present disclosure described above are represented as a series of operations for clarity of description, it is not intended to limit the order in which the steps are performed, and the steps may be performed simultaneously or in different order as necessary. In order to implement the method according to the present disclosure, the described steps may further include other steps, may include remaining steps except for some of the steps, or may include other additional steps except for some of the steps.

[0248] The various implementations of the present disclosure are not a list of all possible combinations and are intended to describe representative aspects of the present disclosure, and the matters described in the various implementations may be applied independently or in combination of two or more.

[0249] In addition, various implementations of the present disclosure may be implemented in hardware, firmware, software, or a combination thereof. In the case of implementing the present disclosure by hardware, the present disclosure can be implemented with application specific integrated circuits (ASICs), Digital signal processors (DSPs), digital signal processing devices (DSPDs), programmable logic devices (PLDs), field programmable gate arrays (FPGAs), general processors, controllers, microcontrollers, microprocessors, etc.

Claims

1.An air mobility apparatus, comprising:an actuator comprising a first propulsion assembly configured to provide lift force and thrust force to the air mobility apparatus;a power source unit comprising (i) a first battery that is configured to generate power and store the generated power, and (ii) a second battery that is rechargeable, the power source unit being configured to output power that is determined based on a flight mode, the flight mode comprising a high-power mode and a low-power mode; andat least one processor configured to control the actuator and the power source unit,wherein the least one processor is configured to:in response to a failure state of the first battery, supply power from the second battery to the first propulsion assembly, the power being determined based on the flight mode, andin response to a failure state of the second battery, supply power from the first battery to the first propulsion assembly, wherein the power supplied from the first battery is greater than or equal to a first power output from the first battery in a normal state of the first battery.2.The air mobility apparatus of claim 1, wherein the least one processor is configured to:based on the normal state of the first battery, supply the first power from the first battery to the first propulsion assembly, the first power being set based on a baseline power of the first propulsion assembly operating in the low-power mode; andbased on the high-power mode and a normal state of the second battery, supply a second power from the second battery to the first propulsion assembly, the second power being set based on a difference between an output power of the first propulsion assembly in the high-power mode and the baseline power of the first propulsion assembly.3.The air mobility apparatus of claim 1, wherein the least one processor is configured to, in response to the failure state of the first battery in the high-power mode, determine the power to be supplied from the second battery second battery based on at least one of (i) an output power of the first propulsion assembly in the high-power mode, (ii) a state of the second battery, or (iii) a flight state of the air mobility apparatus.4.The air mobility apparatus of claim 1, wherein the least one processor is configured to, in response to the failure state of the first battery in the low-power mode, determine the power to be supplied from the second battery based on at least one of (i) the first power output from the first battery in the normal state of the first battery, (ii) a state of the second battery, or (iii) a flight state of the air mobility apparatus.5.The air mobility apparatus of claim 1, wherein the second battery is one of a plurality of second batteries of the power source unit, andwherein the least one processor is configured to, in response to the failure state of one of the plurality of second batteries during the high-power mode, provide power from another second battery in the normal state to the first propulsion assembly, wherein the power to be supplied from the another second battery is determined based on at least one of (i) a difference between an output power of the first propulsion assembly in the high-power mode and a baseline power of the first propulsion assembly, (ii) a state of the another second battery in the normal state, or (iii) a flight state of the air mobility apparatus.6.The air mobility apparatus of claim 5, wherein the actuator further comprises a second propulsion assembly,wherein one of the plurality of second batteries in the normal state is configured to supply power to the first propulsion assembly based on the flight mode,wherein another of the second battery in the normal state is configured to supply power to the second propulsion assembly based on the flight mode, andwherein the least one processor is configured to, in response to the failure state of one of the plurality of second batteries, supply power from one or more second batteries in the normal state to both of the first propulsion assembly and the second propulsion assembly, the power supplied to the second propulsion assembly being determined based on (i) an output power of the second propulsion assembly in the flight mode, (ii) a state of charge of the second battery in the normal state, and (iii) the flight state.7.The air mobility apparatus of claim 6, wherein the least one processor is configured to, in response to the failure state of one of the plurality of second batteries during the high-power mode, supply power from the first battery in the normal state to both of the first propulsion assembly and the second propulsion assembly.8.The air mobility apparatus of claim 6, wherein the first propulsion assembly comprises a tiltable rotor, and the second propulsion assembly comprises a fixed rotor.9.The air mobility apparatus of claim 6, further comprising a plurality of converters that are connected to the plurality of second batteries, respectively, each of the plurality of converters being configured to control an electrical output of one of the plurality of second batteries,wherein the plurality of converters comprise:a primary device configured to control an output voltage of one of the plurality of second batteries, anda secondary device configured to control an output current of another of the plurality of second batteries, andwherein the at least one processor is configured to, in response to the failure state of the second battery connected to the primary device, set one of the plurality of converters connected to the second battery in the normal state as the primary device.10.The air mobility apparatus of claim 1, wherein the flight mode further comprises an emergency landing mode, andwherein the at least one processor is configured to:in the failure state of one of the first battery or the second battery, determine the emergency landing mode based on (i) a battery state of the first battery or the second battery that is operating in the normal state and (ii) a flight state of the air mobility apparatus;in response to the failure state of the first battery, supply power corresponding to the emergency landing mode from the second battery to the first propulsion assembly; andin response to the failure state of the second battery, supply power corresponding to the emergency landing mode from the first battery to the first propulsion assembly.11.The air mobility apparatus of claim 1, wherein the actuator further comprises a second propulsion assembly,wherein the power source unit further comprises a third battery that is rechargeable and has different characteristics from the second battery, andwherein the at least one processor is configured to:based on the first battery, the second battery, and the third battery being in a normal state, supply power from the first battery to the first propulsion assembly;control supply of power from the second battery to the first propulsion assembly and the second propulsion assembly based on the flight mode;control supply of power from the third battery to the second propulsion assembly based on the flight mode; andin response to the failure state of the first battery, supply power from each of the second battery and the third battery to the first propulsion assembly based on at least one of (i) an output power of the first propulsion assembly in the flight mode, (ii) a state of charge of the second and third batteries, or (iii) a flight state of the air mobility apparatus.12.The air mobility apparatus of claim 11, wherein the at least one processor is further configured to disable the supply of power from the second battery to the second propulsion assembly in response to the failure state of the first battery.13.The air mobility apparatus of claim 12, wherein the at least one processor is configured to determine the power supplied from the third battery to the second propulsion assembly based on an output power of the second propulsion assembly in the flight mode.14.The air mobility apparatus of claim 11, wherein the at least one processor is configured to, in response to the failure state of the second battery, supply power from each of the first battery and the third battery to the first propulsion assembly based on at least one of (i) the output power of the first propulsion assembly according to the flight mode, (ii) a state of charge of each of the first battery and the third battery, or (iii) the flight state.15.The air mobility apparatus of claim 14, wherein the at least one processor is configured to, in response to the failure state of the second battery, supply power from the first battery only to the first propulsion assembly.16.The air mobility apparatus of claim 14, wherein the at least one processor is configured to, in response to the failure state of the second battery, supply power from the third battery to the second propulsion assembly based on (i) an output power of the second propulsion assembly in the flight mode, (ii) the states of charge of the first battery and the third battery, and (iii) the flight state.17.The air mobility apparatus of claim 11, wherein the at least one processor is configured to, in response to a failure state of the third battery, supply power from each of the first battery and the second battery to the first propulsion assembly based on at least one of (i) the output power of the first propulsion, (ii) an output power of the second propulsion assembly in the flight mode, (iii) a state of charge of each of the first battery and the second battery module, or (iv) the flight state.18.The air mobility apparatus of claim 17, wherein the at least one processor is configured to, in response to the failure state of the third battery, supply power from each of the first battery and the second battery to the second propulsion assembly based on (i) the output power of the second propulsion assembly, (ii) the states of charge of the first battery and the second battery, and (iii) the flight state.19.The air mobility apparatus of claim 11, wherein the at least one processor is configured to set power of the third battery operating in the normal state based on a baseline power of the first propulsion assembly.20.The air mobility apparatus of claim 19, wherein the at least one processor is configured to determine the power of the second battery supplied to the second propulsion assembly in the high-power mode based on a difference between (i) an output power of the second propulsion assembly in the high-power mode and (ii) the baseline power of the first propulsion assembly.

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