Inverter circuit and electric propulsion system for EVTOL aircraft

Abstract

In one embodiment, an electric propulsion system includes: an electric motor configured to drive one or more propellers of an aircraft; a capacitor configured to stabilize a direct current (DC) bus voltage; a first inverter circuit coupled to the capacitor and configured to convert the DC bus voltage to an alternating current (AC) voltage in response to a first pulse width modulation (PWM) vector to drive a first set of stator windings of the electric motor; and a second inverter circuit coupled to the capacitor and configured to convert the DC bus voltage to an AC voltage to drive a second set of stator windings of the electric motor in response to a second PWM vector. The first PWM vector and the second PWM vector are substantially equal and opposite vectors.

Description

[0001] Cross - reference to related applications

[0002] This disclosure claims priority to U.S. Patent Application No. 18 / 363,535, filed on August 1, 2023, entitled "INVERTER CIRCUITS AND ELECTRICAL PROPULSION SYSTEMS FOR EVTOL AIRCRAFT", which in turn claims priority to U.S. Provisional Application No. 63 / 378,536, filed on October 6, 2022, entitled "Tilt Rotor Systems and Methods for eVTOL Aircraft" and U.S. Provisional Application No. 63 / 378,680, filed on October 7, 2022, entitled "Systems and Methods for Improved Propulsion Systems for eVTOL Aircraft". For all purposes, the contents of these applications are incorporated herein by reference in their entirety. FIELD OF THE DISCLOSURE

[0003] This disclosure generally relates to the field of powered aerial vehicles. More specifically but not limited thereto, this disclosure relates to innovations in tilt-rotor aircraft using electric propulsion systems. Certain aspects of this disclosure generally relate to improvements in electric propulsion systems for tilt-rotor aircraft. Other aspects of this disclosure generally relate to improvements in power inverters that can be used in other types of aircraft but provide specific advantages in aerial vehicles. SUMMARY OF THE DISCLOSURE

[0004] Embodiments of this disclosure provide a propulsion system for an aircraft. The electric propulsion system may include: an electric motor configured to drive one or more propellers of the aircraft; a capacitor configured to stabilize a direct current (DC) bus voltage; a first inverter circuit coupled to the capacitor and configured to convert the DC bus voltage on a first bus of the first inverter circuit to an alternating current (AC) voltage based on a first pulse width modulation (PWM) vector to drive a first set of stator windings of the electric motor; and a second inverter circuit coupled to the capacitor and configured to convert the DC bus voltage on a second bus of the second inverter circuit to an AC voltage based on a second PWM vector to drive a second set of stator windings of the electric motor. The first PWM vector and the second PWM vector are substantially equal and opposite vectors.

[0005] Embodiments of the present disclosure provide a method for controlling a propulsion system for an aircraft. The method may include: stabilizing a direct current (DC) bus voltage through a capacitor; converting the DC bus voltage to an alternating current (AC) voltage by a first inverter circuit coupled to the capacitor according to a first pulse width modulation (PWM) vector to drive the first set of stator windings of the electric motor; and converting the DC bus voltage to an AC voltage by a second inverter circuit coupled to the capacitor in response to a second PWM vector to drive the second set of stator windings of the electric motor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors; and driving one or more propellers of the aircraft by the electric motor. Other embodiments may include corresponding integrated circuits, computer systems, devices, and computer programs recorded on one or more computer storage devices, each computer program configured to perform the operations of the method.

[0006] Embodiments of the present disclosure provide an inverter circuit. The inverter circuit may include: a capacitor configured to stabilize a direct current (DC) bus voltage; a plurality of switches forming a plurality of phase legs, wherein at least one of the phase legs may include an upper switch disposed between a positive terminal of the capacitor and an AC output terminal of the phase leg and a lower switch disposed between a negative terminal of the capacitor and the AC output terminal of the phase leg; and a first discharge circuit coupled in parallel to the capacitor and configured to provide a first discharge path for discharging energy stored in the capacitor. In response to the DC bus voltage being below a threshold in a fault condition, control the plurality of switches to short-circuit the capacitor.

[0007] Embodiments of the present disclosure provide a method for controlling an inverter circuit. The method may include: detecting whether a fault has occurred on one of a plurality of switches in the inverter circuit, disconnecting the inverter circuit from a power source, providing a first discharge path by a first discharge circuit in response to detecting a single-phase short circuit fault to discharge the DC bus voltage across the capacitor of the inverter circuit after the inverter circuit is disconnected from the power source, and controlling the plurality of switches in the inverter circuit to short-circuit the capacitor in response to the bus voltage being below a first threshold. Other embodiments may include corresponding integrated circuits, computer systems, devices, and computer programs recorded on one or more computer storage devices, each computer program configured to perform the operations of the method.

[0008] Embodiments of the present disclosure provide an inverter circuit. The inverter circuit may include: a capacitor configured to stabilize a direct current (DC) bus voltage; a plurality of switches forming a plurality of phase legs, wherein at least one of the phase legs includes an upper switch disposed between a positive terminal of the capacitor and an AC output terminal of the phase leg and a lower switch disposed between a negative terminal of the capacitor and the AC output terminal of the phase leg; a first discharge circuit coupled in parallel to the capacitor and configured to provide a first discharge path for discharging energy stored in the capacitor; and a second discharge circuit coupled in parallel to the capacitor and configured to provide a second discharge path for releasing energy stored in the capacitor in response to the DC bus voltage being lower than a threshold in a fault condition.

[0009] Embodiments of the present disclosure provide a method for controlling an inverter circuit. The method may include: detecting whether a fault has occurred on one of a plurality of switches in the inverter circuit; in response to detecting a single-phase short circuit fault, disconnecting the inverter circuit from a power source; and after disconnecting the inverter circuit from the power source, discharging a bus voltage across a capacitor of the inverter circuit by: in response to confirming that the inverter circuit is disconnected from the power source, using a first discharge circuit to provide a first discharge path; and in response to the bus voltage being lower than a threshold, using a second discharge circuit to provide a second discharge path in parallel with the first discharge path. Other embodiments may include corresponding integrated circuits, computer systems, devices, and computer programs recorded on one or more computer storage devices, each computer program being configured to perform the operations of the method. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] Figure 1 is a diagram of a perspective view of an exemplary VTOL aircraft consistent with some embodiments of the present disclosure.

[0011] Figure 2 is another diagram of a perspective view of an exemplary VTOL aircraft in an alternative configuration consistent with some embodiments of the present disclosure.

[0012] Figure 3 is a diagram of a top plan view of an exemplary VTOL aircraft consistent with some embodiments of the present disclosure.

[0013] Figure 4 is a schematic diagram showing an exemplary propeller rotation of a VTOL aircraft consistent with some embodiments of the present disclosure.

[0014] Figure 5Is a schematic diagram showing an exemplary power connection in a VTOL aircraft consistent with some embodiments of the present disclosure.

[0015] Figure 6 Is a block diagram showing an exemplary architecture and design of an electric propulsion unit of a VTOL aircraft consistent with some embodiments of the present disclosure.

[0016] Figure 7 Is a schematic diagram showing an exemplary tilt electric propulsion system of a VTOL aircraft consistent with some embodiments of the present disclosure.

[0017] Figures 8A - 8C Is an illustration of an exemplary tilt electric propulsion system of a VTOL aircraft consistent with some embodiments of the present disclosure.

[0018] Figure 9 Is a schematic diagram showing an exemplary lift electric propulsion system of a VTOL aircraft consistent with some embodiments of the present disclosure.

[0019] Figures 10A - 10B Is an illustration of an exemplary lift electric propulsion system of a VTOL aircraft consistent with some embodiments of the present disclosure.

[0020] Figure 11 Is a diagram showing a part of an electric propulsion system for a vertical takeoff and landing (VTOL) aircraft consistent with some embodiments of the present disclosure.

[0021] Figure 12A Is a diagram showing an exemplary control of Figure 11 The PWM vectors of the inverter circuit in the electric propulsion system.

[0022] Figure 12B Is a diagram showing the three-phase voltages output by the inverter circuit in one cycle consistent with some embodiments of the present disclosure.

[0023] Figure 13A Is a diagram showing an inverter circuit for motor control in an electric propulsion system consistent with some embodiments of the present disclosure.

[0024] Figure 13B Is a diagram showing another inverter circuit for motor control in an electric propulsion system consistent with some embodiments of the present disclosure.

[0025] Figure 14 Is a diagram showing an exemplary method for controlling Figure 13A Or Figure 13B The inverter circuit of.

[0026] Figure 15is a graph showing the bus voltage of the HV DC bus during a discharge cycle relative to time, in accordance with some embodiments of the present disclosure.

[0027] Figures 16A - 16D is a graph showing the EMI noise in a dual-inverter system or a single-inverter system, in accordance with some embodiments of the present disclosure. Detailed Description

[0028] The following disclosure provides different embodiments or examples for implementing different features of the provided subject matter. Specific examples of components and arrangements are described below to simplify the present disclosure. Of course, these are merely examples and not restrictive. In addition, the present disclosure may repeat reference numerals and / or letters in various examples. This repetition is for the purpose of simplicity and clarity and does not in itself indicate a relationship between the various embodiments and / or configurations discussed.

[0029] The terms used in this specification generally have their ordinary meanings in the art and in the particular context in which each term is used. The examples used in this specification (including examples of any terms discussed herein) are merely illustrative and in no way limit the scope and meaning of the present disclosure or any example terms. Similarly, the present disclosure is not limited to the various embodiments given in this specification.

[0030] Although terms such as "first", "second", etc. may be used herein to describe various elements, these elements should not be limited by these terms. These terms are used to distinguish one element from another. For example, without departing from the scope of the embodiments, a first element may be referred to as a second element, and similarly, a second element may be referred to as a first element. As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items.

[0031] Further, spatial relative terms such as "beneath", "below", "under", "above", "on top" etc. may be used herein to facilitate the description of the relationship of one element or feature depicted in the drawings to another or other elements or one or more features. In addition to the orientation depicted in the drawings, the spatial relative terms are also intended to cover different orientations of the device during use or operation. The device may be oriented in other ways (rotated 90 degrees or in other orientations), and the spatial relative descriptors used herein may be interpreted accordingly.

[0032] In this document, the term "coupled" may also be referred to as "electrically coupled", and the term "connected" may be referred to as "electrically connected". "Coupled" and "connected" may also be used to indicate that two or more elements cooperate or interact with each other.

[0033] The present disclosure relates to components of an electric vertical takeoff and landing (eVTOL) aircraft primarily used in unconventional aircraft. For example, the eVTOL aircraft of the present disclosure may be intended for frequent (e.g., more than 50 flights per weekday), short-duration flights (e.g., less than 100 miles per flight) over, into, and out of populated areas. The aircraft may be intended to carry 4 to 6 passengers or commuters who desire a low-noise and low-vibration experience. Accordingly, it may be required that their components be configured and designed to withstand frequent use without wearing out, that they generate less heat and vibration, and that the aircraft include mechanisms for effectively controlling and managing the heat or vibration generated by the components. Additionally, it may be expected that several of these aircraft operate close to each other over congested metropolitan areas. Accordingly, it may be required that their components be configured and designed to generate a low level of noise both inside and outside the aircraft, and be configured and designed to have various safety and backup mechanisms. For example, for safety reasons, it may be required that the aircraft be propelled by a distributed propulsion system, avoid the risk of single-point failure, and that they be able to take off and land conventionally on a runway. Moreover, it may be required that the aircraft be able to take off vertically and land vertically in a relatively restricted space (e.g., a vertical takeoff and landing airport, a parking lot, or a driveway) compared to a conventional airport runway, while transporting approximately 4 to 6 passengers or commuters with luggage. These usage requirements may impose design constraints on the aircraft size, weight, operational efficiency (e.g., drag, energy usage), which may affect the design and configuration of the aircraft components.

[0034] The disclosed embodiments provide new and improved configurations of aircraft components not observed in conventional aircraft, and / or identify design criteria for components different from those of conventional aircraft. This alternative configuration and design criteria combination addresses the drawbacks and challenges of conventional components, resulting in the embodiments disclosed herein for various configurations and designs of eVTOL aircraft components.

[0035] In some embodiments, the eVTOL aircraft of the present disclosure can be designed to be capable of both vertical and conventional takeoff and landing, where a distributed electric propulsion system enables vertical flight, forward flight, and transition. Thrust can be generated by supplying high-voltage electrical power to electric motors of the distributed electric propulsion system, each of which can convert the high-voltage electrical power into mechanical shaft power to rotate a propeller. Embodiments disclosed herein can relate to optimizing the energy density of the electric propulsion system. Embodiments can include electric motors connected to an on-board electrical power source, which can include a device capable of storing energy, such as a battery or capacitor, or can include one or more systems for harnessing or generating electrical power, such as a fuel-powered generator or a solar panel array. Some disclosed embodiments provide weight reduction and space reduction of components in the aircraft, thereby improving aircraft efficiency and performance. In view of concerns about safety in passenger transportation, the disclosed embodiments implement new and improved safety protocols and system redundancy in the event of a failure to minimize any single point of failure in the aircraft propulsion system. Some disclosed embodiments also provide new and improved methods to meet aviation and transportation laws and regulations. For example, the Federal Aviation Administration of the United States enforces federal laws and regulations that require safety components, such as fire barriers, to be near engines that use more than a threshold amount of oil or other flammable materials.

[0036] In a preferred embodiment, the distributed electric propulsion system can include twelve electric motors, which can be mounted on booms at the front and rear of the main wing of the aircraft. The front electric motors can be capable of tilting between a horizontal orientation position (e.g., to generate forward thrust) and a vertical orientation position (e.g., to generate vertical lift) during flight. In terms of the direction of propeller rotation, the front electric motors can have a clockwise type or a counterclockwise type. The rear electric motors can be fixed in a vertical orientation position (e.g., to generate vertical lift). In terms of the direction of propeller rotation, these rear electric motors can also have a clockwise type or a counterclockwise type. In some embodiments, the aircraft can have various combinations of front and rear electric motors. For example, the aircraft can have six front electric motors and six rear electric motors, four front electric motors and four rear electric motors, or any other combination of front and rear engines, including embodiments where the number of front and rear electric motors is not equal. In some embodiments, the aircraft can have four front propellers and four rear propellers, where at least four of these propellers include tiltable propellers.

[0037] In a preferred embodiment, for vertical takeoff and landing (VTOL) missions, the front electric engine and the rear electric engine can provide vertical thrust during takeoff and landing. During the flight phase when the aircraft is in the forward flight mode, the front electric engine can provide horizontal thrust, while the propellers of the rear electric engine can be retracted in a fixed position to minimize drag. The rear electric engine can be actively retracted using position monitoring. The transition from vertical flight to horizontal flight and vice versa can be achieved via a tilt propeller subsystem. The tilt propeller subsystem can redirect thrust between the primary vertical direction during the vertical flight mode and the primary horizontal direction during the forward flight mode. A variable pitch mechanism can change the blade collective angle of the propeller hub assembly of the front electric engine for operation during the hover phase, transition phase, and cruise phase.

[0038] In some embodiments, for conventional takeoff and landing (CTOL) missions, the front electric engine can provide horizontal thrust for wing-borne takeoff, cruise, and landing. In some embodiments, the rear electric engine may not be used to generate thrust during CTOL missions and the rear propeller can be retracted in place.

[0039] In some embodiments, the electric engine can be housed or attached to a strut of the aircraft and includes a motor, an inverter, and a gearbox. In some embodiments, the motor, inverter, and gearbox can be joined such that they share a central axis. In some embodiments, torque from the motor can be sent out from the propeller of the propulsion system and sent to the gearbox. In some embodiments, the gearbox can provide gear reduction and then send the torque back through the spindle to a bearing inside the motor and back to the propeller. In some embodiments, the inverter can be mounted at the rear of the gearbox such that the spindle does not travel through the inverter when outputting torque to the propeller. In some embodiments, the motor, gearbox, and inverter can be joined such that a coolant, such as oil, can be used to maintain the motor, inverter, and / or gearbox while sharing a common heat exchanger. In some embodiments, the amount of oil used for lubricating and cooling the electric engine can vary and include amounts less than one quart, two quarts, three quarts, or any other measured amount of oil.

[0040] In some embodiments, the tilt propeller system may include a linear or rotary actuator to change the orientation of the propulsion system during operation. In some embodiments, the pitch of the propulsion system may be changed according to the orientation of the propulsion system. In some embodiments, the rotary actuator may include a motor, an inverter, and a gearbox. In some embodiments, the gearbox may include various types of gears that engage to provide a gear reduction capable of orienting the propulsion system. In some embodiments, the tilt propeller system may include a redundant configuration such that there are multiple motors, inverters, and gearboxes and they are engaged using gears. In some embodiments, a configuration utilizing multiple motors, gearboxes, and inverters may allow a failed portion of the redundant configuration to be driven by the motors, inverters, and gearboxes of another portion of the configuration. In some embodiments, the gearbox configuration may also allow the tilt propeller system to maintain the propulsion system orientation with the help of additional power provided by the system or without the additional power provided by the system.

[0041] In some embodiments, a electric propulsion system as described herein can generate thrust by supplying high voltage (HV) electrical power to an electric engine, which in turn converts the HV electrical power into mechanical shaft power for rotating a propeller. As described above, an aircraft as described herein may have multiple electric engines mounted on booms at the front and rear of the wings. The amount of thrust generated by each electric engine can be controlled by a torque command sent by a flight control system (FCS) to each electric engine via a digital communication interface. Embodiments may include a front electric engine and may be capable of changing its orientation or tilting. Additional embodiments include a front engine that may be of a clockwise (CW) type or a counterclockwise (CCW) type. The front electric engine propulsion subsystem may consist of a multi-blade adjustable pitch propeller and a variable pitch subsystem.

[0042] In some embodiments, the aircraft may include a tail engine or elevator that may be of a clockwise (CW) type or a counterclockwise (CCW) type. Additional embodiments may include a tail electric engine utilizing a multi-blade fixed pitch propeller.

[0043] As described herein, the orientation and use of an electric propulsion system can be varied throughout the operation of an aircraft. In some embodiments, during vertical takeoff and landing, both a front propulsion system and a tail propulsion system can provide vertical thrust during takeoff and landing. During the flight phase when the aircraft is in forward flight mode, the front propulsion system can provide horizontal thrust while the propellers of the tail propulsion system can be retracted in a fixed position to minimize drag. The tail electric propulsion system can be actively retracted using position monitoring. Some embodiments can include a transition from vertical flight to horizontal flight and vice versa. In some embodiments, the transition can be achieved via a tilt propeller system (TPS). The TPS redirects thrust between a primarily vertical direction during vertical flight mode and a primarily horizontal direction during forward flight mode. Additional embodiments can include a variable pitch mechanism that can vary the total blade angle of the front propulsion system propeller hub assembly for operation during hover, cruise, and transition phases. Some embodiments can include a conventional takeoff and landing (CTOL) configuration such that the tilter provides horizontal thrust for wing-borne takeoff, cruise, and landing. The tail electric engine is not used to generate thrust during CTOL missions and the tail propeller is retracted in a suitable position.

[0044] In some embodiments, an electric engine as described herein can have design features to mitigate and prevent non - contained fires, such as using a non - hazardous amount of combustible fluid contained in both the tilt engine and the lift engine, having no nominal ignition source within the electric engine, having an engine over - temperature operating limit that can be 50 °C or more below the auto - ignition temperature of the combustible fluid, over - temperature detection and protection, over - voltage detection and protection, and over - current detection and protection. In some embodiments, the design features of the electric engine can be such that it is not a designated fire - proof area. In some embodiments, the combustible fluid can include oil and the non - hazardous amount can be less than one quart, or two quarts, or three quarts, or four quarts, or five quarts, or ten quarts, as determined based on factors such as the size of the aircraft, the number of propellers, or the payload.

[0045] As disclosed herein, an electric engine can include an inverter and a motor; or an inverter, a gearbox, and a motor in various configurations, such as the representative configurations described herein. For example, an electric engine can include an electric motor, a gearbox, and an inverter, all sharing the same central axis. Additionally, the central axis can be configured along the axis of the output shaft leading to the aircraft's propeller. In this exemplary configuration, the motor, gearbox, and inverter will all share the output shaft as the central axis and will be oriented circularly around the output shaft. Additional embodiments can include a motor, a gearbox, and an inverter mounted together in sequence, or a configuration where some components (such as the motor and gearbox) are mounted together and another component (such as the inverter) is located elsewhere but uses a wiring system to connect to the electric engine.

[0046] As described above, the electric engine for an aircraft as described herein may include some or all of a motor, an inverter, and a gearbox. Various configurations may include an inverter and a motor such that the output shaft of the motor directly provides speed and torque to the propeller shaft. Additional embodiments of the electric engine may include a motor, an inverter, and a gearbox, where the output of the motor may travel through a gearbox connected to the output shaft of the propeller; a motor, an inverter, and a gearbox, where the output from the motor travels away from the propeller through a gearbox, where the output shaft of the propeller travels through the gearbox and the motor and returns to the propeller. As described herein, the electric engine may contemplate any combination or orientation of some or all of the motor, the inverter, and the gearbox. Additionally, each configuration or orientation of the electric engine as disclosed herein may include cooling via air cooling, coolant liquid, or a mixture of both.

[0047] For example, a configuration of the electric engine may include a motor and an inverter, where the motor is located between the propeller of the aircraft and the inverter. Additionally, the motor may include a gearbox. Further, the inverter may share the same central axis with the motor, where the inverter may be located in a housing cantilevered from the rear of the motor and may be air-cooled. It should be recognized that this inverter orientation may not be the optimal configuration in terms of the housing required to achieve such a cantilever orientation. Additionally, the motor in such a configuration utilizing air cooling may include potting material, and the air fins assisting in motor cooling may result in an even greater increase in system mass.

[0048] Some embodiments may include an electric engine where the inverter module may be mounted external to the motor housing. Additional embodiments may include an electric engine where the inverter may be mounted on top of the electric motor such that the air-cooling fins of the inverter are below the propeller. Further embodiments may include: the inverter mounted to the rear of the motor where the air-cooling fins face radially outward; the inverter mounted to the front of the motor where the air-cooling fins face radially outward; the inverter mounted to the motor where the inverter is liquid-cooled, for example, by oil; or any other position of the inverter relative to the motor.

[0049] Embodiments of the electric motor may include a stator housing, a wound stator assembly, a rotor, various bearings, and any additional components such that it facilitates the transfer of the speed and torque generated by the motor to the propeller.

[0050] It should be understood that an electric engine can generate heat during operation and may include a thermal management system to ensure that the components of the electric engine do not fail during operation. In some embodiments, a coolant can be used and circulated through various components of the engine, such as an inverter, a gearbox, or a motor, either through some components or through all components of the engine, to help manage the heat present in the engine. Additional embodiments may include using an air cooling method to cool the electric engine, or using a mixture of coolant and air to manage the heat generated during operation in the electric engine. In some embodiments, the coolant used can also be the same liquid that is used as a lubricant throughout the inverter, gearbox, or motor. For example, liquid or air can be used to cool the inverter, gearbox, and motor, or a combination of air and liquid cooling can be used, such as using air cooling to cool the motor and liquid cooling in the inverter and gearbox, or any other combination of air and liquid cooling on the inverter, gearbox, and motor or even a subset of these components.

[0051] In some embodiments, oil can be used as a lubricant throughout the electric engine and can also be used as a coolant fluid to help manage the heat generated by the engine during operation. Further for this example, different amounts of oil can be used to act as a lubricant and coolant fluid in the electric engine, with or without air cooling assistance, such as less than one quart, less than two quarts, or any other amount of oil required to lubricate and cool the electric engine. As has been disclosed herein, an electric engine can have different primary functions, such as being used only for lifting and landing and thus being used only in one orientation, or being used during all phases of flight, such as lifting, landing, and flying in between. An engine used during all phases of flight can experience various orientations throughout the flight and may include more lubricant and coolant than an engine used only in one orientation. Therefore, all engines on an aircraft may not contain the same amount of lubricant and coolant. For example, a lift and landing engine may only require less than one quart of oil, while an engine operating during all phases of flight may require more than one quart of oil. It should be understood that the example embodiments mentioned herein are representative and do not prescribe a limit on the amount of lubricant and coolant that can be used in an electric engine.

[0052] It should be understood that by using oil to not only lubricate but also cool the electric engine, rather than using another coolant, this will add additional oil to the system, but this additional oil will eliminate the conventional components available for cooling such an electric engine. For example, if the electric engine is cooled by another liquid such as ethylene glycol, the engine may include separate heat exchangers for both the lubricant fluid and the coolant fluid. Thus, in embodiments where a single fluid (such as oil) is used for lubrication and cooling, there will be an increase in oil, but only one heat exchanger will be required, so the mass of the entire system can be reduced due to the use of fewer heat exchangers and potentially no other components, and there may be a more attractive drag profile. Additionally, due to the reduction in mass and the benefits of cooling the engine with one substance rather than relying on air cooling which may have issues traveling throughout the engine, using one substance to lubricate and cool the engine can improve the efficiency of the system.

[0053] Additional embodiments of the electric engine can include various components to ensure monitoring and prevention of any combustible fluid from entering certain sections of the electric engine. Some embodiments may include an electric engine having a wet zone housing, which can be defined by a gearbox, a motor, and / or a heat exchanger. In some embodiments, the electric engine can have up to four liters or more of air in contact with the engine oil within the motor-gearbox housing. For example, based on factors such as the size, number of propellers, or payload of the aircraft, the electric engine can have up to five liters, or six liters, or eight liters, or ten liters, or twenty liters of air in contact with the engine oil within the motor-gearbox housing. Embodiments of the motor-gearbox housing can use a breather to equalize the internal pressure and the external pressure. Embodiments of the breather can include a breather that protrudes above nearby design features to prevent external fluid from inadvertently entering. Additional embodiments can include a breather having a screen and a tortuous entry path to prevent external debris from entering. Embodiments can include viewing windows present on both the inclined electric engine and the lift electric engine to check during maintenance that the oil is not overfilled or underfilled.

[0054] Additional embodiments of the electric engine may include active protection features in the front and rear electric engines, such as monitoring the internal temperature of the entire engine operation, including oil temperature, stator winding kit, inverter bulk capacitor, power module, control board power module, control board control processor, control board monitoring processor, internal hot spots, and various other locations throughout the engine. Embodiments may include over-temperature limits considering known failure temperatures and operating limits related to the auto-ignition temperature of the fluid. Some embodiments may include a high-voltage power system that may have a fuse at the high-voltage battery terminals, which can irreversibly and quickly disconnect the engine electrical connection to mitigate over-current events. This over-current protection can be activated when the current consumption of the electric engine is greater than the over-current operation. Thus, in some embodiments, a fault condition that causes over-current may only result in transient overheating, arcing, or spark faults. Some embodiments may include a fire threat characterization test ignition source, which can be selected to be a more severe ignition source than a short circuit that occurs in the electric engine and is disconnected by the engine fuse. In some embodiments, the inverter will detect AC over-current and isolate the faulty phase and / or continuously monitor the input DC voltage, and will apply protection actions to keep the voltage below the over-voltage operation limit.

[0055] A. Exemplary Electric Aircraft Features

[0056] Figure 1 is an illustration of a perspective view of an exemplary VTOL aircraft that is consistent with the disclosed embodiments. Figure 2 is another illustration of a perspective view of an exemplary VTOL aircraft in an alternative configuration that is consistent with the embodiments of the present disclosure. Figure 1 and Figure 2 show VTOL aircraft 100, 200 that are consistent with the embodiments of the present disclosure in a cruise configuration and a vertical takeoff, landing, and hover configuration (also referred to herein as the "lift" configuration), respectively. Elements corresponding to Figure 1 and Figure 2 may have similar reference numerals and refer to similar elements of the aircraft 100, 200. The aircraft 100, 200 may include fuselages 102, 202, wings 104, 204 mounted to the fuselages 102, 202, and one or more rear stabilizers 106, 206 mounted to the rear of the fuselages 102, 202. A plurality of lift propellers 112, 212 may be mounted to the wings 104, 204 and may be configured to provide lift for vertical takeoff, landing, and hovering. A plurality of tilt propellers 114, 214 may be mounted to the wings 104, 204 and may be tiltable between a lift configuration and a cruise configuration, in which the plurality of tilt propellers provide a portion of the lift required for vertical takeoff, landing, and hovering, as Figure 2As shown, in the cruise configuration, the plurality of tilt propellers provide forward thrust to the aircraft 100 for horizontal flight, as Figure 1 shown. As used herein, a tilt propeller lift configuration refers to any tilt propeller orientation in which tilt propeller thrust primarily provides lift to the aircraft, and a tilt propeller cruise configuration refers to any tilt propeller orientation in which tilt propeller thrust primarily provides forward thrust to the aircraft.

[0057] In some embodiments, the lift propellers 112, 212 may be configured to provide only lift, where all horizontal propulsive forces are provided by the tilt propellers. Thus, the lift propellers 112, 212 may be configured to have a fixed position and may generate thrust only during the takeoff, landing, and hover phases of flight. Meanwhile, the tilt propellers 114, 214 may be tilted upward into a lift configuration in which the thrust from the propellers 114, 214 is directed downward to provide additional lift.

[0058] For forward flight, the tilt propellers 114, 214 may be tilted from their lift configuration to their cruise configuration. In other words, the orientation of the tilt propellers 114, 214 may change from an orientation in which tilt propeller thrust is directed downward (to provide lift during vertical takeoff, landing, and hover) to an orientation in which tilt propeller thrust is directed backward (to provide forward thrust to the aircraft 100, 200). The tilt propeller assemblies for a particular electric engine may be tilted about a rotational axis defined by the mounting points connecting the struts and the electric engine. When the aircraft 100, 200 are in full forward flight, lift may be provided entirely by the wings 104, 204. Meanwhile, in the cruise configuration, the lift propellers 112, 212 may be shut off. The blades 120, 220 of the lift propellers 112, 212 may be held in a low-drag position for aircraft cruise. In some embodiments, the lift propellers 112, 212 may each have two blades 120, 220, and the two blades may be locked in a minimum-drag position for cruise, in which one blade is directly in front of the other blade, as Figure 1 shown. In some embodiments, the lift propellers 112, 212 have more than two blades. In some embodiments, the tilt propellers 114, 214 may include more blades 116, 216 than the lift propellers 112, 212. For example, as Figure 1 and Figure 2 shown, the lift propellers 112, 212 may each include, for example, two blades, while the tilt propellers 114, 214 may each include more blades, such as the five blades shown. In some embodiments, each of the tilt propellers 114, 214 may have from 2 to 5 blades, and possibly more, depending on the design considerations and requirements of the aircraft.

[0059] In some embodiments, the aircraft may include a single wing 104, 204 on each side of the fuselage 102, 202 (or a single wing extending across the entire aircraft). At least a portion of the lift propellers 112, 212 may be located behind the wings 104, 204, and at least a portion of the tilt propellers 114, 214 may be located in front of the wings 104, 204. In some embodiments, all of the lift propellers 112, 212 may be located behind the wings 104, 204, and all of the tilt propellers 114, 214 may be located in front of the wings 104, 204. According to some embodiments, all of the lift propellers 112, 212 and tilt propellers 114, 214 may be mounted to the wings - that is, neither the lift propellers nor the tilt propellers are mounted to the fuselage. In some embodiments, the lift propellers 112, 212 may all be located behind the wings 104, 204, and the tilt propellers 114, 214 may all be located in front of the wings 104, 204. According to some embodiments, all of the lift propellers 112, 212 and tilt propellers 114, 214 may be positioned inside the ends of the wings 104, 204.

[0060] In some embodiments, the lift propellers 112, 212 and tilt propellers 114, 214 may be mounted to the wings 104, 204 via struts 122, 222. The struts 122, 222 may be mounted below the wings 104, 204, on top of the wings, and / or may be integrated into the wing profile. In some embodiments, the lift propellers 112, 212 and tilt propellers 114, 214 may be mounted directly to the wings 104, 204. In some embodiments, each strut 122, 222 may be mounted with a lift propeller 112, 212 and a tilt propeller 114, 214. The lift propellers 112, 212 may be mounted at the rear end of the struts 122, 222, and the tilt propellers 114, 214 may be mounted at the front end of the struts 122, 222. In some embodiments, the lift propellers 112, 212 may be mounted in fixed positions on the struts 122, 222. In some embodiments, the tilt propellers 114, 214 may be mounted to the front end of the struts 122, 222 via hinges. The tilt propellers 114, 214 may be mounted to the struts 122, 222 such that the tilt propellers 114, 214 are aligned with the body of the struts 122, 222 when in their cruise configuration, thereby forming a continuous extension of the front end of the struts 122, 222 that minimizes the drag during forward flight.

[0061] In some embodiments, the aircraft 100, 200 may include, for example, one wing on each side of the fuselage 102, 202 or a single wing extending across the aircraft. According to some embodiments, at least one wing 104, 204 is a high wing mounted to the upper side of the fuselage 102, 202. According to some embodiments, the wing includes control surfaces, such as flaps and / or ailerons. According to some embodiments, the wings 104, 204 may have been designed with a profile that reduces drag during forward flight. In some embodiments, the wing tip profile may be curved and / or tapered to minimize drag.

[0062] In some embodiments, the rear stabilizers 106, 206 include control surfaces, such as one or more rudders, one or more elevators, and / or one or more combined rudder-elevators. The wing can have any suitable design. In some embodiments, the wing has a tapered leading edge.

[0063] In some embodiments, the lift propellers 112, 212 or tilt propellers 114, 214 may be deflected relative to at least one other lift propeller 112, 212 or tilt propeller 114, 214. As used herein, deflection refers to the relative orientation of the axis of rotation of the lift propeller / tilt propeller about a line parallel to the forward-backward direction, similar to the roll degree of freedom of an aircraft. Deflection of the lift propeller and / or tilt propeller can be achieved by orienting the plane of rotation of the lift propeller / tilt propeller disk (blades plus the hub to which the blades are mounted) so as not to intersect critical parts of the aircraft (such body regions where people may be located, critical flight control systems, batteries, adjacent propellers, etc.) or other propeller disks, thereby assisting in minimizing damage caused by propeller burst and providing enhanced yaw control during flight.

[0064] Figure 3 is an illustration of a top plan view of an exemplary VTOL aircraft consistent with embodiments of the present disclosure. The aircraft 300 shown in the figure may be respectively at Figure 1 and Figure 2Top plan views of the aircraft 100, 200 shown. As discussed herein, the aircraft 300 can include twelve electric propulsion systems distributed across the aircraft 300. In some embodiments, the distribution of the electric propulsion systems can include six front electric propulsion systems 314 and six rear electric propulsion systems 312 mounted on booms at the front and rear of the main wing 304 of the aircraft 300. In some embodiments, the length of the boom 324 from the wing 304 to the rear end of the lift propeller can include a similar boom 324 rear end length across multiple rear ends of the boom. In some embodiments, the rear end length of the boom can vary across the exemplary six rear ends of the boom. For example, each rear end of the boom 324 can include a different length from the wing 304 to the lift propeller, or a subgroup of the rear ends of the boom can be similar in length. In some embodiments, the front end of the boom 322 can include various lengths from the wing 304 to the tilt propeller across the front end of the boom. For example, as Figure 3 shown, the front end length of the boom 322 from the tilt propeller closest to the fuselage to the wing 304 can include a greater length than the front end length of the boom 322 from the wing 304 to the tilt propeller furthest from the fuselage. Some embodiments can include a front end of the boom that has a similar length across the exemplary six front ends of the boom, or any other distribution of the front end length of the boom from the wing 304 to the tilt propeller. Some embodiments can include an aircraft 300 having eight electric propulsion systems, where the eight electric propulsion systems have four front electric propulsion systems 314 and four rear electric propulsion systems 312, or any other distribution of front and rear electric propulsion systems, including embodiments where the number of front electric propulsion systems 314 is less than or greater than the number of rear electric propulsion systems 312. Additionally, Figure 3 depicts an exemplary embodiment of a VTOL aircraft 300 having front propellers in a horizontal orientation for horizontal flight and tail propeller blades 320 in a stowed position for the forward flight phase.

[0065] As disclosed herein, the front electric propulsion systems and the rear electric propulsion systems can be of the clockwise (CW) type or the counterclockwise (CCW) type. Some embodiments can include various front electric propulsion systems having a mixture of both CW type and CCW type. In some embodiments, the rear electric propulsion systems can have a mixture of CW type and CCW type systems among the rear electric propulsion systems.

[0066] Figure 4 is a schematic diagram showing an exemplary propeller rotation of a VTOL aircraft consistent with the disclosed embodiments. The aircraft 400 shown in the figure can be respectively at Figure 1 , Figure 2 and Figure 3Top plan views of the aircraft 100, 200, and 300 shown. The aircraft 400 may include six front electric propulsion systems, where three of these front electric propulsion systems have a CW type 424, and the remaining three front electric propulsion systems have a CCW type. In some embodiments, three rear electric propulsion systems may have a CCW type 428, while the remaining three rear electric propulsion systems may have a CW type 430. Some embodiments may include an aircraft 400 having four front electric propulsion systems and four rear electric propulsion systems, each of the four front electric propulsion systems and the four rear electric propulsion systems having two CW types and two CCW types. In some embodiments, the propellers may rotate in opposite directions relative to adjacent propellers to eliminate torque steering experienced by the fuselage or wings of the aircraft due to the rotation of the propellers. In some embodiments, the difference in the direction of rotation may be achieved using the direction of rotation of the engines. In other embodiments, the engines may all rotate in the same direction, and a gear arrangement may be used to achieve different propeller rotation directions.

[0067] Some embodiments may include an aircraft 400 having front electric propulsion systems and rear electric propulsion systems, where the amounts of CW type 424 and CCW type 426 are not equal among the front electric propulsion systems, among the rear electric propulsion systems, or among the front electric propulsion systems and the rear electric propulsion systems.

[0068] Figure 5FIG. 0 is a schematic diagram showing an exemplary power connection in a VTOL aircraft consistent with the disclosed embodiments. The VTOL aircraft may have various power systems connected to diagonally opposed electric propulsion systems. In some embodiments, the power system may include a high-voltage power system. Some embodiments may include a high-voltage power system connected to an electric engine via a high-voltage channel. In some embodiments, the aircraft 500 may include six power systems 526, 528, 530, 532, 534, and 536, which include batteries stored within the wings 570 of the aircraft 500. In some embodiments, the aircraft 500 may include six front electric propulsion systems having six electric engines 502, 504, 506, 508, 510, and 512 and six rear electric propulsion systems having six electric engines 514, 516, 518, 520, 522, and 524. In some embodiments, the batteries may be connected to diagonally opposed electric engines. In this configuration, the first power system 526 may supply power to the electric engine 502 via the power connection channel 538 and to the electric engine 524 via the power connection channel 540. In some embodiments, the first power system 526 may also be paired with the fourth power system 532 via the power connection channel 542, and the power connection channel 542 has a fuse to prevent excessive current from flowing through the power systems 526 and 532. Further for this embodiment, the VTOL aircraft 500 may include a second power system 528 paired with the fifth power system 534 via a power connection channel 548 having a fuse, and may supply power to the electric engines 510 and 516 via the power connection channels 544 and 546, respectively. In some embodiments, the third power system 530 may be paired with the sixth power system 536 via a power connection channel 554 having a fuse, and may supply power to the electric engines 506 and 520 via the power connection channels 550 and 552, respectively. The fourth power system 532 may also supply power to the electric engines 508 and 518 via the power connection channels 556 and 558, respectively. The fifth power system 534 may also supply power to the electric engines 504 and 522 via the power connection channels 560 and 562, respectively. The sixth power system 536 may also supply power to the electric engines 512 and 514 via the power connection channels 564 and 566, respectively.

[0069] As disclosed herein, an electric propulsion system may include an electric engine connected via a high-voltage channel or a power connection channel to a high-voltage power system (such as a battery) located within an aircraft. Some embodiments may include various batteries stored within the wings of the aircraft, the wings of the aircraft having a high-voltage channel that travels through the aircraft (including the wings and booms) to the electric propulsion system. In some embodiments, multiple high-voltage power systems may be used to create an electric propulsion system having multiple high-voltage power sources to avoid the risk of single-point failure. In some embodiments, the aircraft may include multiple electric propulsion systems, the multiple electric propulsion systems being wired in a certain pattern to various batteries or power sources stored throughout the aircraft. It should be appreciated that such a configuration may be beneficial in avoiding the risk of single-point failure, where the failure of one battery or power source may cause a portion of the aircraft to be unable to maintain the thrust required to continue flying or to make a controlled landing. For example, if a VTOL has two front electric propulsion systems and two rear propulsion systems, the front electric propulsion system and the rear propulsion system on opposite sides of the VTOL aircraft may be connected to the same high-voltage power system. In such a configuration, if one high-voltage power system fails, the front electric propulsion system and the rear propulsion system on opposite sides of the VTOL aircraft will remain operational and may provide a more balanced flight or landing compared to the front electric propulsion system and the rear propulsion system that have failed on the same side of the VTOL aircraft. Some embodiments may include four front electric propulsion systems and four rear electric propulsion systems, where diagonally opposite electric engines are connected to a common battery or power source. Some embodiments may include various configurations of electric engines electrically connected to a high-voltage power system such that the risk of single-point failure is avoided in the event of a power source failure and the flight phase may continue during the occurrence of the failure, or the aircraft may perform an alternative flight phase in response to the failure.

[0070] As discussed above, an electric propulsion system may include an electric engine that provides mechanical shaft power to a propeller assembly to generate thrust. In some embodiments, the electric engine of the electric propulsion system may include a high-voltage power system that supplies high-voltage power to the electric engine and / or a low-voltage system that supplies low-voltage DC power to the electric engine. Some embodiments may include an electric engine that digitally communicates with a flight control system ("FCS") including a flight control computer ("FCC"), the flight control system being capable of sending signals to and receiving signals from the electric engine, the signals including commands and response data or status. Some embodiments may include an electric engine capable of receiving operating parameters from the FCC and transmitting operating parameters to the FCC, these operating parameters including speed, voltage, current, torque, temperature, vibration, propeller position, and any other values of operating parameters.

[0071] In some embodiments, a flight control system may include a system that is capable of communicating with an electric engine to send and receive analog / digital signals to and from the electric engine and control a device that is capable of redirecting the thrust of a tilt propeller between a primary vertical direction during a vertical flight mode and a primary horizontal direction during a forward flight mode. In some embodiments, this system may be referred to as a tilt propeller system (“TPS”) and is capable of communicating with and orienting additional features of an electric propulsion system.

[0072] Figure 6 A block diagram showing an exemplary architecture and design of an electric propulsion unit 600 consistent with the disclosed embodiments is shown. In some embodiments, an electric propulsion system 602 may include an electric engine subsystem 604 that may supply torque via an axial propeller subsystem 606 to generate thrust for the electric propulsion system 602. Some embodiments may include an electric engine subsystem 604 that receives low voltage DC (LV DC) power from a low voltage system (LVS) 608. Some embodiments may include an electric engine subsystem 604 that receives high voltage (HV) power from a high voltage power system (HVPS) 610 that includes at least one battery or another device capable of storing energy. In some embodiments, the high voltage power system may include more than one battery or another device capable of storing energy that supplies high voltage power to the electric engine subsystem 604. It should be appreciated that this configuration may be advantageous because there is no risk of a single point of failure where a single battery failure causes the electric propulsion system 602 to fail.

[0073] Some embodiments may include an electric propulsion system 602 that includes an electric engine subsystem 604 that receives signals from and transmits signals to a flight control system 612. In some embodiments, the flight control system 612 may include a flight control computer that is capable of using Controller Area Network (“CAN”) data bus signals to send commands to and receive status and data from the electric engine subsystem 604. It should be understood that although CAN data bus signals are used between the flight control computer and the electric engine, some embodiments may include any form of communication capable of sending and receiving data from the flight control computer to the electric engine. In some embodiments, the flight control system 612 may also include a tilt propeller system (“TPS”) 614 that is capable of sending analog discrete data to and receiving analog discrete data from the electric engine subsystem 604 of the tilt propeller. The tilt propeller system 614 may include devices that are capable of transferring operating parameters to the electric engine subsystem 604 and defining the orientation of the propeller subsystem 606 to use mechanical components, such as gearbox assemblies, linear actuators, and any other configured components, during various phases of flight to change the orientation of the propeller subsystem 606 to redirect the thrust of the tilt propeller.

[0074] As discussed throughout, an exemplary VTOL aircraft may include various types of electric propulsion systems that include tilt propellers and lift propellers that include a front electric engine capable of tilting during various phases of flight and a rear electric engine that remains in one orientation and may be active only during certain phases of flight (i.e., takeoff, landing, and hover).

[0075] Figure 7 is a schematic diagram showing an exemplary tilt electric propulsion system of a VTOL aircraft consistent with the disclosed embodiments. The tiltable electric propulsion system 700 may include an electric engine assembly 702 aligned along an axis 724 that is connected to an output shaft 738 that is mechanically coupled to a propeller assembly 720 that includes a hub, a spinner, and tilt propeller blades. In some embodiments, the electric engine assembly 702 may include a motor and gearbox assembly 704 aligned along the axis 724 and mechanically coupled to the axis. In some embodiments, the motor and gearbox assembly 704 may include an electric motor assembly that includes a stator 706 and a rotor 708. As Figure 7As shown and present in some embodiments, the stator 706 may include a plurality of stator windings connected to an inverter 716. In such a configuration, the stator 706 may include one or more redundancies such that in the event of a failure of a set of windings, power will still be transmitted to the stator 706 via one or more remaining windings, such that the electric power engine assembly 702 maintains power and continues to generate thrust at the propeller assembly 720.

[0076] In some embodiments, the motor and gearbox assembly 704 may include a gearbox 710 aligned along the shaft 724 to provide a gear reduction between the torque of the shaft 724 from the electric power engine assembly including the stator 706 and the rotor 708 and the output shaft 738. The torque applied to the output shaft 738 may be transmitted to the propeller assembly 720. Some embodiments may include a gearbox 710 containing an oil pump. In such embodiments, the oil pump may drive oil to circulate throughout the motor and gearbox assembly 704 at a speed equal to the rotation of the output shaft 738 to cool and lubricate the gearbox and electric motor components. In some embodiments, the oil pump may drive the circulation of oil at a speed greater than or less than the rotation of the output shaft 738. Some embodiments of the motor and gearbox assembly 704 may include a propeller position sensor 712 present within the housing, which may detect the magnetic field generated by the electric power engine assembly to determine the propeller position. Further embodiments may include a propeller position sensor 712 that is powered by the inverter 716 and sends the collected data to the inverter 716.

[0077] In some embodiments, the electric power engine assembly 702 may further include an inverter assembly 714 aligned substantially along the shaft 724. The inverter assembly 714 may include an inverter 716 and an inverter power source 740. The inverter power source 740 may receive low-voltage DC power from a low-voltage system 734 located outside the electric power engine assembly 702. The inverter power source 740 may receive low-voltage DC power that has been converted to low-voltage DC power via a DC-DC converter 742 from a high-voltage power system 732 located outside the electric power engine assembly 702. The inverter 716 may supply high-voltage alternating current (AC) to the stator 706 of the electric power engine assembly within the motor and gearbox assembly 704 via at least one three-phase winding. The inverter assembly 714 may include an inverter 716 that may receive flight control data from a flight control computing subsystem 736.

[0078] In some embodiments, the motor and gearbox 704 can be located between the inverter assembly 714 and the propeller assembly 720. Some embodiments can also include a separator plate 744 coupled to the motor and gearbox assembly 704 and the inverter assembly 714. The separator plate 744 can create an enclosed environment for the upper portion of the motor and gearbox assembly 704 via a bell end cap assembly and an enclosed environment for the lower portion of the inverter assembly 714 via a hot plate. In some embodiments, the separator plate 744 can function as an integral mounting bracket for supporting the heat exchanger 718. The heat exchanger 718 can include, for example, a folded fin or other type of heat exchanger. In some embodiments, the electric propulsion system 700 can circulate oil or other coolant throughout the electric engine assembly 702, the motor and gearbox assembly 704, or the inverter assembly 714 to transfer heat generated from the components to the oil or other coolant liquid. The heated oil or other coolant liquid can be circulated through the heat exchanger 718 to transfer heat to the air stream 722 passing through the fins of the heat exchanger.

[0079] In some embodiments, the electric engine assembly 702 can be mounted or coupled to the pylon structure 726 of the aircraft. The variable pitch mechanism 730 can be mechanically coupled to the propeller assembly 720. In some embodiments, the variable pitch mechanism can be adjacent to the electric engine assembly 702. In some embodiments, the variable pitch mechanism 730 can be coupled to the variable pitch mechanism 730 such that it can be remotely mounted within the pylon, wing, or fuselage of the aircraft. In some embodiments, the variable pitch mechanism 730 can be included within or travel adjacent to the shaft 724 to the shaft or components of the propeller assembly 720. The variable pitch mechanism 730 can change the total blade angle of the propeller assembly of the front electric engine during the hover, transition, and cruise phases according to the operational requirements. Some embodiments can include an electric engine assembly 702 mechanically coupled to the tilt propeller subsystem 728, which can redirect the thrust between the predominantly vertical direction during vertical flight mode and the predominantly horizontal direction during forward flight mode. In some embodiments, the tilt propeller subsystem can be adjacent to the variable pitch mechanism 730. Some embodiments can include a tilt propeller subsystem 728, which includes various components located at various positions. For example, the components of the tilt propeller subsystem can be coupled to the electric engine assembly 702, and other components can be coupled to the variable pitch mechanism 730. These various components of the tilt propeller subsystem 728 can work together to redirect the thrust of the tiltable electric propulsion system 700.

[0080] Figures 8A - 8C is an illustration of an exemplary tilt electric propulsion system of a VTOL aircraft consistent with the disclosed embodiments. Figures 8A - 8CHave like reference numerals and refer to like elements of tiltable electric propulsion systems 800A, 800B, and 800C. Thus, similar design considerations and configurations may be considered throughout the embodiments.

[0081] Figure 8A and 8B show side profiles and perspective views, respectively, of tiltable electric propulsion systems 800A, 800B in a cruise configuration integrated into booms 812A, 812B in accordance with the present disclosure. The tiltable propeller electric propulsion systems 800A, 800B may include electric engine assemblies 802A, 802B housed within booms 812A, 812B of a VTOL aircraft. In some embodiments, the cruise configuration may include electric engine assemblies 802A, 802B disposed within booms 812A, 812B. As described herein, the electric engine assemblies 802A, 802B may include an electric motor assembly, a gearbox assembly, an inverter assembly having power connection channels 810A, 810B, and heat exchangers 804A, 804B. The electric engine assemblies 802A, 802B may be mechanically coupled to propulsion assemblies 808A, 808B, which include shaft flange assemblies 806A, 806B, a spinner, and propeller blades.

[0082] Figure 8C shows a top view along spinner 808C of tiltable electric propulsion system 800C in a lift configuration integrated into boom 812B in accordance with the present disclosure. As Figure 8C shown, the tiltable electric propulsion system 800C in the lift configuration may include electric engine assemblies 802A, 802B that are disposed external to boom 812C and that change their orientation relative to boom 812C.

[0083] As discussed herein, a lift electric propulsion system may be configured to provide thrust in one direction and may not provide thrust during all phases of flight. For example, the lift system may provide thrust during takeoff, landing, and hover, but may not provide thrust during cruise.

[0084] Figure 9It is a schematic diagram showing an exemplary lift electric propulsion system of a VTOL aircraft consistent with the disclosed embodiments. The lift electric propulsion system 900 can be mounted or coupled to the boom structure 924 of the aircraft. The lift electric propulsion system 900 can include an electric engine assembly 902 aligned along an axis 940, the axis being connected to an output shaft 932, the output shaft being mechanically coupled to a propeller assembly 920, the propeller assembly including a hub and tilting propeller blades. In some embodiments, the electric engine assembly 902 can include a motor and gearbox assembly housing 904 aligned along the axis 940 and mechanically coupled to the axis. In some embodiments, the motor and gearbox assembly housing 904 can include an electric motor assembly, the electric motor assembly including a stator 906 and a rotor 908. The stator 906 can include a plurality of stator windings connected to an inverter 916. In such a configuration, the stator 906 can include one or more redundancy and backup measures to avoid single points of failure in such cases. For example, the stator 906 can include a plurality of windings such that if a winding fails, power can continue to be transferred to the stator 906 via the remaining windings, thereby allowing the electric engine assembly 902 to maintain power and continue to generate thrust at the propeller assembly 920.

[0085] In some embodiments, the motor and gearbox assembly housing 904 can include a gearbox 910 aligned along the axis 940 to provide a gear reduction between the torque of the axis 932 of the electric engine assembly including the stator 906 and the rotor 908 and the output shaft 932. The torque applied to the output shaft 932 can be transferred to the propeller assembly 920. Some embodiments can include a gearbox 910 that includes a fluid pump for circulating a cooling fluid and / or a lubricating fluid. In the illustrated embodiment, the fluid pump is an oil pump. In such an embodiment, the oil pump can drive oil to circulate throughout the motor and gearbox assembly housing 904 at a speed equal to the rotation of the output shaft 932 to cool and lubricate the gearbox and electric motor components. Some embodiments of the motor and gearbox assembly housing 904 can include a propeller position sensor 912 present within the housing, the propeller position sensor can detect the magnetic field generated by the electric engine assembly to determine the propeller position. Further embodiments can include a propeller position sensor 912 that is powered by the inverter 916 and sends the collected data to the inverter 916, the collected data can be transferred to the flight control computing system 930 together with other flight control data.

[0086] In some embodiments, the electric engine assembly 902 may further include an inverter assembly housing 914 aligned along an axis shared with the axis of the shaft. The inverter assembly housing 914 may include an inverter 916 and an inverter power source 934. The inverter power source 934 may receive low-voltage DC power from a low-voltage system 928 located outside the electric engine assembly 902. The inverter power source 934 may receive low-voltage DC power that has been converted to low-voltage DC power via a DC-DC converter 936 from a high-voltage power system 926 located outside the electric engine assembly 902. The inverter 916 may supply high-voltage alternating current to the stator 906 of the electric engine assembly located within the motor and gearbox assembly housing 904 via at least one three-phase winding. The inverter assembly 914 may include the inverter 916, which may send data to and receive data from the flight control computing subsystem 930.

[0087] In some embodiments, the motor and gearbox housing 904 may be located between the inverter assembly housing 914 and the propeller assembly 920. Some embodiments may further include a separator plate 938 coupled to the motor and gearbox assembly housing 904 and the inverter assembly housing 914. The separator plate 938 may create a sealed environment for the upper portion of the motor and gearbox assembly housing 904 via a bell end cap assembly and may create a sealed environment for the lower portion of the inverter assembly housing 914 via a heat plate. In some embodiments, the separator plate 938 may act as an integral mounting bracket for supporting the heat exchanger 918. The heat exchanger 918 may include, for example, folded fins or other types of heat exchangers. In some embodiments, the electric propulsion system 900 may circulate oil or other coolant fluid throughout the electric engine assembly 902, the motor and gearbox assembly 904, or the inverter assembly 914 to transfer heat generated from the components to the oil or other coolant liquid. The heated oil or other coolant liquid may be circulated through the heat exchanger 918 to transfer heat to the air stream 922 passing through the fins of the heat exchanger.

[0088] In some embodiments, the tiltable electric propulsion system and the lift electric propulsion system may have similar components. This can be advantageous for many design considerations present within a VTOL aircraft. For example, from a manufacturability perspective, different types of electric propulsion systems with similar components can be beneficial in terms of manufacturing efficiency. Additionally, having similar components can be beneficial in terms of risk management because similar components have similar points of failure, and these points of failure can be well explored and designed when comparing systems with similar components to systems with different components and configurations.

[0089] Although a tiltable electric propulsion system can have additional and in some embodiments different components compared to a lift electric propulsion system, it should be understood that in some embodiments, a tiltable electric propulsion system and a lift electric propulsion system can have the same component configuration. For example, in some embodiments, a tiltable electric propulsion system and a lift electric propulsion system can include the same components, while the lift electric propulsion system can be coupled to a strut, wing, or fuselage of an aircraft such that it may not be able to provide thrust in as many directions as a tiltable electric propulsion system.

[0090] Figures 10A - 10B is an illustration of an exemplary lift electric propulsion system of a VTOL aircraft consistent with the disclosed embodiments. Figure 10A and Figure 10B have like reference numerals and refer to like elements of lift electric propulsion systems 1000A and 1000B. Thus, similar design considerations and configurations can be considered throughout the embodiments.

[0091] Figure 10A Shows a side profile of lift electric propulsion system 1000A in a lift configuration integrated into strut 1010A consistent with the present disclosure. Lift electric propulsion system 1000A can include an electric engine assembly 1002A housed within strut 1010A of a VTOL aircraft. In some embodiments, the lift configuration can include electric engine assembly 1002A vertically disposed within strut 1010A. As described herein, electric engine assembly 1002A can include an electric motor assembly, a gearbox assembly, an inverter assembly having a power connection channel 1008A, and a heat exchanger 1004A. Electric engine assembly 1002A can be mechanically coupled to propulsion assembly 1006A, which includes a shaft flange assembly and propeller blades.

[0092] Figure 10B Shows a top view of lift electric propulsion system 1000B in a lift configuration integrated into strut 1010B consistent with the present disclosure.

[0093] Some embodiments of the disclosed electric engines can generate heat during operation and may include a thermal management system to ensure that components of the electric engine do not fail during operation. In some embodiments, a coolant can be used and circulated through various components of the engine (such as an inverter, gearbox, or motor), through some components or through all components of the engine, to help manage the heat present in the engine. Some embodiments may include using an air cooling method to cool the electric engine, or using a mixture of coolant and air to manage the heat generated during operation in the electric engine. In some embodiments, the coolant used can also be the same liquid that is used as a lubricant throughout the inverter, gearbox, or motor. For example, a liquid or air or a mixture of air and liquid cooling can be used to cool components of the electric engine. As another example, air cooling can be used to cool the motor, while liquid cooling can be used to cool the inverter and gearbox. It should be understood that hybrid cooling can be used for any combination of electric engine components or within each component.

[0094] In some embodiments, oil can be used as a lubricant throughout the electric engine and can also be used as a coolant fluid to help manage the heat generated by the engine during operation. Further for this example, different amounts of oil can be used to act as a lubricant and coolant fluid in the electric engine, with or without air cooling assistance, such as less than or equal to 1 quart, 1.5 quarts, 2 quarts, 2.5 quarts, 3 quarts, 5 quarts, or any other amount of oil required to lubricate and cool the electric engine. In some embodiments, the amount of oil or liquid to be used in the system related to cooling can be determined based on the amount of thermal mass required to drive heat transfer from components of the electric propulsion system. As has been disclosed herein, an electric engine can have different primary functions, such as being used only for lifting and landing and thus only used in one orientation, or being used during all phases of flight such as lifting, landing, and flying in the air, etc. An engine used during all phases of flight can experience various orientations throughout the flight and may include more lubricant and coolant than an engine used only in one orientation. Therefore, all engines on an aircraft may not contain the same amount of lubricant and coolant. For example, a lift and landing engine may only require less than one quart of oil, while an engine operating during all phases of flight may require more than one quart of oil. In some embodiments, the amount of oil or liquid used for cooling can be an appropriate amount to provide sufficient thermal mass to drive heat transfer from components of the electric propulsion system, regardless of the orientation of the electric propulsion system. The embodiments discussed herein are exemplary, non-limiting, and do not prescribe bounds on the amount of lubricant and coolant that can be used in an electric engine.

[0095] Some embodiments may use oil to lubricate and cool an electric engine. Such embodiments may require an additional volume of oil. In such embodiments, the additional oil may permit removal of conventional components that may be used to cool such an electric engine. For example, if the electric engine is cooled by another fluid such as ethylene glycol, the engine may include separate heat exchangers for both the lubricant fluid and the coolant fluid. Thus, in embodiments that use a single fluid (such as oil) for both lubrication and cooling, there will be an increase in oil, but only one heat exchanger will be required, so the mass of the overall system may be reduced due to the use of fewer heat exchangers and potentially no other components, and there may be a more attractive drag profile. Additionally, due to the reduced mass and the benefits of cooling the engine with a substance rather than relying on air cooling that may have issues traveling throughout the engine, using one substance to lubricate and cool the engine can improve the efficiency of the system.

[0096] Some embodiments of an electric engine may include various components for monitoring a combustible fluid and for preventing combustible materials from entering certain sections of the electric engine. Some embodiments may include an electric engine having a wet area housing that may be defined by a gearbox, a motor, and / or a heat exchanger. In some embodiments, the electric engine may have up to 4 liters or more of air within the motor-gearbox housing that is in contact with engine oil. Embodiments of the motor-gearbox housing may use a breather to equalize internal and external pressures. Embodiments of the breather may include a breather that protrudes above nearby design features to prevent external fluids from inadvertently entering. Some embodiments may include a breather having a screen and a tortuous entry path to prevent external debris from entering. Embodiments may include a viewing window that is present on both the inclined electric engine and the lift electric engine to check for overfilling or underfilling of oil during maintenance.

[0097] Some embodiments of an electric engine may include active protection features in a front electric engine and a tail electric engine, such as monitoring the vibration and internal temperature of the entire engine, such as oil temperature, stator winding kit temperature, inverter bulk capacitor temperature, power module temperature, control board power module temperature, control board control processor temperature, control board monitoring processor temperature, internal hot spot temperature, and various other operating conditions of the entire engine as needed. Such monitoring can be done using various sensors located throughout the electric propulsion system and the aircraft. Embodiments may include vibration limits based on known failure points or resonances of components and over-temperature limits based on known failure temperatures and operating limits related to the autoignition temperature of fluids. In some embodiments, the various sensors for monitoring the operating conditions of the entire engine may report the operating conditions to a flight control system. Some embodiments may include threshold operating values that may be required before the operating values are sent to or flagged by the flight control system. In some embodiments, the flight control system may act in response to detecting an operating condition to reduce the amount of power directed to the electric propulsion system. Some embodiments may include reducing the amount of power to the electric propulsion system to reduce mechanical wear or frictional sparks caused by vibration, and / or reducing power in an attempt to lower the temperature of components present within the electric propulsion system. Additionally, some embodiments may include reducing the power to the electric propulsion system in the case where the detected efficiency of the inverter is less than a target efficiency. In some embodiments, for example, in the case where there are twelve electric propulsion systems within an aircraft, the flight control system may be used to reduce the power to or terminate the power to a single electric propulsion system while increasing the power directed to the remaining electric propulsion systems or a subset thereof to counteract the reduction in lift generated by one electric propulsion system. In some embodiments, the flight control system may establish various thresholds of operating conditions to correspond to a reduction or increase in the power to the electric propulsion system.

[0098] Some embodiments may include a high-voltage power system that may have a fuse at the high-voltage battery terminals, which can quickly and irreversibly disconnect the engine electrical connection to mitigate and avoid overcurrent events. Such overcurrent protection can be activated when the current draw of the electric engine is greater than an overcurrent operation. Thus, in some embodiments, a fault condition that results in an overcurrent may only result in transient overheating, arcing, or spark faults. Some embodiments may include a fire threat characterization test ignition source that may be selected to be a more severe ignition source than a short circuit that occurs within the electric engine and is disconnected by the engine fuse. In some embodiments, the inverter may detect an AC overcurrent and isolate the faulty phase and / or continuously monitor the input DC voltage and will apply protective actions to keep the voltage below an overvoltage operating limit.

[0099] During takeoff, landing, hovering, and cruising, the motors and associated control components of a VTOL aircraft can generate heat. The heat must be dissipated to prevent degradation or damage to the motors, control components, and other components of the VTOL aircraft. For some types of VTOL aircraft, such as electric VTOL (eVTOL) aircraft, thermal control is also important for maintaining optimal energy efficiency of components powered by, for example, batteries.

[0100] Some components may generate high thermal loads only during certain operating periods. For example, some lift propellers can be used only during takeoff, landing, and hovering and can be turned off during cruising. Thus, such lift propellers can generate high thermal loads during takeoff, landing, and hovering and little or no heat during cruising.

[0101] B. Example Inverter Embodiments

[0102] Figure 11 Shows a portion of an electric propulsion system 1100 for a vertical takeoff and landing (VTOL) aircraft consistent with some embodiments of the present disclosure. The electric propulsion system 1100 can provide a dual three-phase system for motor control. As Figure 11 shown, the electric propulsion system 1100 includes a first inverter circuit 1110, a second inverter circuit 1120, an electric motor M1 configured to drive one or more propellers of the VTOL aircraft, and a bus capacitor 1170 configured to stabilize a direct current (DC) bus voltage Vbus. The first inverter circuit 1110 is coupled to the bus capacitor 1170 and is configured to convert the DC bus voltage Vbus on the bus of the first inverter circuit 1110 into an alternating current (AC) voltage in response to a first pulse width modulation (PWM) vector to drive a first set of stator windings of the electric motor M1. The second inverter circuit 1120 is configured to convert the DC bus voltage Vbus on the bus of the second inverter circuit 1120 into an AC voltage in response to a second PWM vector to drive a second set of stator windings of the electric motor M1. In some embodiments, the first PWM vector and the second PWM vector are substantially equal and opposite vectors. For example, the delay between the PWM signals corresponding to the first PWM vector and the second PWM vector can be equal to or less than 0.25%, 0.5%, 1%, or 2% of the switching cycle period. For example, the delay can be within 50 nanoseconds. Thus, the first inverter circuit 1110 is configured to output a first set of three-phase AC voltages (e.g., u1, v1, w1), and the second inverter circuit 1120 is configured to output a second set of three-phase AC voltages (e.g., u2, v2, w2), and the phases of the first set of three-phase AC voltages and the corresponding phases of the second set of three-phase AC voltages are two interleaved phases having a phase shift of substantially 180 degrees (e.g., plus or minus 5 degrees).

[0103] Specifically, in the electric propulsion system 1100, two inverter circuits 1110 and 1120 are electrically coupled to an internal high-voltage DC power bus and configured to provide corresponding three-phase AC voltages u1, v1, and w1 and three-phase AC voltages u2, v2, and w2 to drive a dual three-phase motor M1. Figure 11 The dual-inverter drive system shown in FIG. can improve motor performance and system reliability by increasing the number of phases.

[0104] As shown, inverter circuits 1110 and 1120 are respectively configured to convert the bus voltage Vbus on the high-voltage DC power bus into three-phase AC power to drive motor M1. When inverter circuits 1110 and 1120 convert DC power into AC power, there is a voltage difference between the neutral points of the power source and the load, which is called the common-mode voltage. The common-mode current caused by the common-mode voltage in the inverter may be harmful to the electrical system. Specifically, the common-mode voltage may cause motor failure, premature bearing failure, glitches in control equipment, etc. To reduce common-mode noise, filter components can be installed in the electric propulsion system 1100.

[0105] For example, the electric propulsion system 1100 may include a DC common-mode filter 1130 and AC common-mode chokes 1140 and 1150. The DC common-mode filter 1130 can be coupled to the bus capacitor 1170 and is configured to reduce the common-mode signals on the DC sides of the first inverter circuit 1110 and the second inverter circuit 1120. One or more AC common-mode chokes 1140 and 1150 can be coupled to the AC sides of the first inverter circuit 1110 or the second inverter circuit 1120 to reduce the common-mode signals.

[0106] For example, the DC common-mode filter 1130 can be located between the DC power source 1160 and the bus capacitor 1170 and is formed by a set of DC-side chokes 1132 and a set of DC common-mode filter capacitors 1134, 1136. The DC-side chokes 1132 can be configured such that the positive and negative lines are wound around the same magnetic core. Therefore, the DC-side chokes 1132 and the DC common-mode filter capacitors 1134, 1136 can be configured to reduce the common-mode signals on the DC side. However, larger filters may increase the volume and mass and may incur additional losses.

[0107] In some embodiments, the electric propulsion system 1100 can achieve common-mode voltage cancellation by applying appropriate space vector modulation (SVM) to the inverter circuits 1110 and 1120.

[0108] In various embodiments, the winding arrangements of the motor M1 applied in the electric propulsion system 1100 can be different. For example, the phase difference of the electrical angle between two sets of three-phase windings can be designed to reduce harmonic components. In some embodiments, the first set of stator windings and the second set of stator windings of the motor M1 are offset by substantially 180 degrees (e.g., plus or minus 5 degrees). That is, the motor phasing between the independent windings of each set can be 180 degrees or approximately 180 degrees out of phase.

[0109] In some embodiments, standard center-aligned space vector modulation (SVM) is used to control the first inverter circuit 1110, while reverse center-aligned space vector modulation is used to control the second inverter circuit 1120. In this mode of operation, the common-mode voltage can be eliminated by equal and opposite PWM vectors.

[0110] Figure 12A is a diagram showing the PWM vectors for controlling Figure 11 the first inverter circuit 1110 and the second inverter circuit 1120 in the electric propulsion system 1100 that is consistent with some embodiments of the present disclosure. Figure 12B is a diagram showing the three-phase voltages u1, v1, and w1 output by the first inverter circuit 1110 and the three-phase voltages u2, v2, and w2 output by the second inverter circuit 1120 within one cycle that is consistent with some embodiments of the present disclosure.

[0111] In some embodiments, the space vector modulation (SVM) algorithm is applied to control pulse width modulation (PWM) and to generate an AC voltage from a DC voltage to drive a three-phase motor at varying speeds. As will be understood, different SVM algorithms can have different qualities and computational requirements. As Figure 12A shown, for a three-leg inverter using space vector modulation, there are eight possible switching vectors SV0 - SV7. Figure 12A shows an example PWM vector V INV 1 of the first inverter circuit 1110 and an example PWM vector V INV 2 of the second inverter circuit 1120.

[0112] During operation, the switches within the inverter circuits 1110 and 1120 are controlled such that two switches in the same leg (i.e., the upper switch and the lower switch) are not turned on simultaneously to avoid short-circuiting the DC power supply. This can be achieved by complementary operation of the switches within the same leg. That is, for each output leg, when the upper switch is turned on, the lower switch is turned off, and vice versa. Therefore, the switching vectors SV0 - SV7 include six effective switching vectors SV1 - SV6 and two zero vectors SV0 and SV7.

[0113] As Figure 12AAs shown, the switching vector SV0 = {000} indicates that the upper switches of the three phases U, V, and W are turned off, while the lower switches of the three phases U, V, and W are turned on. The switching vector SV1 = {100} indicates that the upper switch of phase U is turned on, and the upper switches of phases V and W are turned off. The switching vector SV2 = {110} indicates that the upper switches of phases U and V are turned on, and the upper switch of phase W is turned off. The switching vector SV3 = {010} indicates that the upper switch of phase V is turned on, and the upper switches of phases U and W are turned off. The switching vector SV4 = {011} indicates that the upper switches of phases V and W are turned on, and the upper switch of phase U is turned off. The switching vector SV5 = {001} indicates that the upper switch of phase W is turned on, and the upper switches of phases U and V are turned off. The switching vector SV6 = {101} indicates that the upper switches of phases U and W are turned on, and the upper switch of phase V is turned off. The switching vector SV7 = {111} indicates that the upper switches of the three phases U, V, and W are turned on, while the lower switches of the three phases U, V, and W are turned off.

[0114] In some embodiments, the first inverter circuit 1110 is controlled according to a standard center - aligned SVM starting from the switching vector SV0, where the upper switches of the three phases U, V, and W are turned off, and the second inverter circuit 1120 is controlled according to an inverted center - aligned SVM starting from the switching vector SV7, where the upper switches of the three phases U, V, and W are turned on. Thus, as Figure 12B shown, in each stage of the operating cycle, the three - phase voltages u1, v1, w1 output by the first inverter circuit 1110 are respectively complementary to the three - phase voltages u2, v2, w2 output by the second inverter circuit 1120. The resulting PWM vector V INV 1 of the first inverter circuit 1110 and the PWM vector V INV 2 of the second inverter circuit 1120 are equal and opposite PWM vectors. Therefore, the common - mode voltage V INV1-CM generated by the first inverter circuit 1110 and the common - mode voltage V INV2-CM generated by the second inverter circuit 1120 have the same value but opposite signs, and the total common - mode voltage V SYS-CM is zero due to the cancellation of the common - mode voltages.

[0115] In some embodiments, a six-phase machine can thus be implemented to achieve similar functionality to a three-phase machine, but with reduced common-mode noise experienced by the system. In some embodiments, common-mode voltage cancellation can be achieved across any number of phases such that a combination of PWM vectors can cancel or reduce the common-mode noise experienced by the system. In some embodiments, the first inverter circuit 1110 and the second inverter circuit 1120 can be configured to simultaneously draw power from a power source (e.g., from the DC bus capacitor 1170), and the first inverter circuit 1110 and the second inverter circuit 1120 operate in accordance with PWM signals having equal duty cycles but substantially 180 degrees out of phase.

[0116] By implementing common-mode voltage cancellation by the electric propulsion system 1100 having two inverters, the volume and mass of electromagnetic compatibility (EMC) components required in the system can be reduced. In some embodiments, the common-mode voltage cancellation can provide a lower noise level of at least 30 - 40 dB at low frequencies, which has a substantial impact on the filter size. For example, the required size of the magnetic core for the DC side choke can be significantly reduced, which reduces the total weight of the engine.

[0117] Reference Figure 13A , which shows an inverter circuit 1300 for motor control in an electric propulsion system consistent with some embodiments of the present disclosure. The inverter circuit 1300, as a voltage source inverter (VSI), can use FETs S1 - S6 as switches. The FETs S1 - S6 can be selectively turned on or off in response to control signals from a controller in accordance with the SVM algorithm described above to generate a three-phase AC output voltage to drive the motor M1.

[0118] As Figure 13A shown in the embodiments of

[0119] In the inverter circuit 1300, the FETs S1 - S6 can operate independently. During operation, when any one of the FETs S1 - S6 fails, a single - phase short - circuit fault may occur, where one of the bridge arms shorts, resulting in an uncontrolled current through the corresponding phase when the machine is rotating. This type of fault can lead to a fire hazard, create drag on the system, impose very high torque ripples on the system, or any other potential danger. Therefore, it would be advantageous to reduce the torque present in response to a single - phase short - circuit fault in the system operation.

[0120] The inverter circuit 1300 includes a first discharge circuit 1310 across two terminals of the bus capacitor 1320. That is, the first discharge circuit 1310 is coupled in parallel to the bus capacitor 1320. In some embodiments, the first discharge circuit 1310 includes a discharge resistor 1312 and a switch 1314 connected in series to the discharge resistor 1312 to provide a discharge path for releasing the energy stored in the bus capacitor 1320 (i.e., the bulk capacitor) when the switch 1314 closes in response to a corresponding command signal from the controller. Thus, when a single - phase short - circuit occurs, the inverter circuit 1300 can discharge the bus capacitor 1320 by closing the switch 1314 to remove the energy stored in the bus capacitor 1320 at an appropriate timing.

[0121] In some embodiments, in response to the DC bus voltage Vbus being below a threshold in a fault condition, the FETs S1 - S6 can be further controlled to short - circuit the bus capacitor 1320. In other words, during the discharge of the bus capacitor 1320, when the bus voltage Vbus is below a desired safety threshold, bridge short - through can be applied to all the FETs S1 - S6 to short - circuit the HV bus to ensure a safe discharge process and spread the heat generated during a single - phase fault across all the FETs S1 - S6 in the inverter circuit 1300 to ensure safety and reduce damage when a fault occurs.

[0122] In addition, as Figure 13A shown, pyro fuses F1 and F2 are coupled between the inverter circuit 1300 and the DC voltage source Vin (e.g., a battery pack). For example, the pyro fuses F1 and F2 can be a type of fuse configured to be activated by an external source when circuit disconnection and isolation are required. For example, the first pyro fuse F1 can be coupled between the positive terminal of the DC voltage source Vin and the positive terminal of the bus capacitor 1320. The second pyro fuse F2 can be coupled between the negative terminal of the DC voltage source Vin and the negative terminal of the bus capacitor 1320.

[0123] Referring to Figure 13B, which shows another inverter circuit 1300 for motor control in a power propulsion system consistent with some embodiments of the present disclosure. Figure 13B The inverter circuit 1300 is also a voltage source inverter (VSI) that uses FETs S1 - S6 as switches to output a three - phase AC output voltage to drive the motor M1.

[0124] Compared with Figure 13A the inverter circuit 1300 of Figure 13B the inverter circuit 1300 further includes a comparator circuit 1330 and a second discharge circuit 1340 that form another discharge path across the bus capacitor 1320.

[0125] Specifically, in Figure 13B the embodiment of

[0126] when the inverter circuit 1300 discharges the bus capacitor 1320, the comparator circuit 1330 is configured to monitor the bus voltage Vbus across the bus capacitor 1320. In some other embodiments, the comparator circuit 1330 can also be configured to monitor the voltage across any other component within the inverter circuit 1300 to determine whether to release energy through the second discharge circuit 1340 based on the voltage measurement. The second discharge circuit 1340 is configured to provide a fast discharge path for the remaining energy stored in the bus capacitor 1320. In some embodiments, the threshold can be designed to confirm that the power source (e.g., battery) has been disconnected from the HV DC bus.

[0127] When the bus capacitor 1320 discharges through the first discharge circuit 1310, the bus voltage Vbus gradually decreases, and the voltage V1, which is a fixed fraction of the bus voltage Vbus, also decreases. When the voltage V1 is lower than the reference voltage Vref, it is determined that the bus voltage Vbus is lower than a preset threshold, indicating that the power source is disconnected from the HV DC bus. In response to the voltage V1 being lower than the reference voltage Vref, the comparator 1334 is configured to output a corresponding signal (e.g., a logic one signal). The logic circuit 1336 can be an AND gate configured to receive the signal output by the comparator 1334 and the command signal Cmd from the control circuit. When both the command signal Cmd and the signal output by the comparator 1334 are logic ones, the logic circuit 1336 can output a corresponding control signal to turn on the switch 1344 in the second discharge circuit 1340. Thus, the bus capacitor 1320 can discharge through the second discharge circuit 1340 to achieve rapid discharge.

[0128] For example, the fast discharge component 1342 can be a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance value device, or any combination thereof, but the present disclosure is not limited thereto. The fast discharge component 1342 can be any other component capable of achieving rapid discharge to remove the remaining energy on the HV DC bus.

[0129] In other words, according to the above circuit operation, rapid high-voltage discharge of the inverter circuit 1300 can be achieved by monitoring the voltage across components (e.g., bus capacitor) in the inverter circuit 1300 and performing a through-conduction across the FETs S1 - S6 in the inverter circuit 1300. In some embodiments, after initiating the first discharge phase to start voltage discharge, the voltage across the bus capacitor is monitored using a comparator circuit until a threshold is reached. In response to detecting that the monitored voltage drops to the threshold, the second discharge phase can be carried out in parallel. The second discharge phase can utilize a discharge component (e.g., the fast discharge component 1342) such as a TVS or an MOV to discharge the remaining voltage accordingly. In some other embodiments, the above threshold voltage detection may not be the only condition for initiating the second discharge phase. For example, in addition to detecting that the monitored voltage drops to the threshold, a corresponding command (e.g., the command signal Cmd) for initiating the second discharge phase can be generated and sent to one or more components (e.g., the logic circuit 1336) to initiate the second discharge phase. In various embodiments, the specific threshold voltage can depend on battery characteristics, resistor characteristics, discharge time length, or any other components involved in the circuit system.

[0130] Figure 14 Illustrated is for controlling in accordance with some embodiments of the present disclosure Figure 13A or Figure 13BExample flowchart of method 1400 for inverter circuit 1300 to mitigate single-phase short circuit faults. As Figure 14 shown, method 1400 may include steps 1410, 1420, 1430, 1440, and 1450.

[0131] In step 1410, it is detected whether a fault occurs on one of a plurality of switches (e.g., FETs S1 - S6) in inverter circuit 1300. In some embodiments, inverter circuit 1300 is configured to confirm whether a single-phase short circuit fault occurs on any one of FETs S1 - S6. For example, inverter circuit 1300 may include voltage or current sensing circuits or components to detect single-phase short circuit faults based on voltage or current signals within inverter circuit 1300.

[0132] In step 1420, in response to detecting a single-phase short circuit fault, a three-phase short circuit is applied by controlling low-side switches (e.g., Figure 13B FETs S2, S4, and S6 therein) or high-side switches (e.g., Figure 13B FETs S1, S3, and S5 therein). In some embodiments, in response to confirming that a single-phase short circuit fault has occurred, in step 1420, inverter circuit 1300 is configured to apply a three-phase short circuit on the normal side of inverter circuit 1300. For example, when a low-side FET (e.g., S2, S4, or S6) is damaged resulting in a single-phase short circuit, the high-side FETs (e.g., S1, S3, and S5) can be controlled to apply a three-phase short circuit, and vice versa.

[0133] After applying the three-phase short circuit, in step 1430, in response to detecting a single-phase short circuit fault, inverter circuit 1300 is disconnected from the power source. In some embodiments, inverter circuit 1300 is configured to instruct the battery management system (BMS) in the system to enable a protection mechanism to disconnect the HV circuit from the faulty inverter circuit 1300. In some embodiments, the battery management system is housed within a high-voltage junction box (HVJB) and is configured to monitor voltage, temperature, current, and isolation resistance, and control battery pack contactors and pyrotechnic fuses to prevent fault conditions for safe operation.

[0134] In some embodiments, step 1430 includes steps 1432, 1434, and 1436. For example, in step 1432, an instruction can be sent from inverter circuit 1300 to the BMS. Then, in step 1434, the BMS can send one or more command signals to one or more pyrotechnic fuse drivers to activate one or more pyrotechnic fuses, thereby disconnecting inverter circuit 1300 from the power source.

[0135] In other words, in response to an instruction or command from the inverter circuit 1300, the BMS is configured to generate a pyrotechnic event to activate one or both of the pyrotechnic fuses F1 and F2, thereby disconnecting the HV DC voltage source Vin from the faulty inverter circuit 1300. Additionally, in step 1436, the BMS may further send an acknowledgement signal to the inverter circuit 1300 after one or more pyrotechnic fuses have been activated, confirming that the pyrotechnic fuses have been fired.

[0136] That is, when a short - circuit event occurs, the BMS is used to activate the pyrotechnic fuses F1 and / or F2. In some embodiments, the BMS may also send a command signal to the corresponding pyrotechnic fuse driver to activate the pyrotechnic fuses, so as to prevent over - current and electrically isolate the battery pack from the connected inverter circuit 1300 when other types of faults occur. Then, the inverter circuit 1300 may confirm that the HV DC voltage source Vin (e.g., the battery pack) is disconnected from the inverter circuit 1300 in response to receiving the acknowledgement signal.

[0137] In step 1440, after the inverter circuit 1300 is disconnected from the power source, the bus voltage Vbus across the bus capacitor 1320 of the inverter circuit 1300 is discharged. In some embodiments, after receiving the acknowledgement signal, the inverter circuit 1300 is configured to remove the energy stored in the bus capacitor 1320. For example, the inverter circuit 1300 may close the switch 1314, thereby forming a discharge path to discharge the energy using the discharge resistor 1312. Thus, the bus voltage Vbus across the bus capacitor 1320 gradually decreases during the discharge process. In some embodiments, in step 1440, the inverter circuit 1300 may perform a rapid discharge by using a plurality of parallel discharge circuits.

[0138] For example, in step 1442, in response to confirming that the inverter circuit 1300 is disconnected from the power source, the first switch (e.g., the switch 1314 in Figure 13B the first discharge circuit 1310) in the first discharge circuit (e.g., Figure 13B is closed to provide a first discharge path. In step 1444, the bus voltage Vbus across the bus capacitor 1320 is monitored by a comparator circuit (e.g., the comparator circuit 1330 in Figure 13B ). In step 1446, in response to the bus voltage Vbus being lower than a second threshold, the second switch (e.g., the switch 1344 in Figure 13B the second discharge circuit 1340) in the second discharge circuit (e.g., Figure 13B is closed to provide a second discharge path in parallel with the first discharge path.

[0139] In step 1450, in response to the bus voltage Vbus being lower than a first threshold, a plurality of switches (e.g., Figure 13B FETs S1 - S6 in Figure 13B ) in the inverter circuit 1300 are closed to short - circuit the bus capacitor 1320. In some embodiments, the inverter circuit 1300 is configured to detect the bus voltage Vbus across the bus capacitor 1320, and when the bus voltage Vbus is lower than a specific threshold voltage (e.g., about 50 V), the inverter circuit 1300 can apply a bridge - arm shoot - through to short - circuit all six FETs S1 - S6, thereby short - circuiting the high - voltage bus. The threshold voltage can be designed based on actual needs. For example, the range of the threshold voltage can be about 40 V - 60 V. Since the shoot - through configuration is enabled after the bus voltage Vbus drops below a certain level during the fast - discharge process, the resulting current will be within the safe current limit and will not damage the components (e.g., FETs S1 - S6) in the inverter circuit 1300. After the high - voltage bus is short - circuited, the motor M1 is stopped accordingly.

[0140] By the operation of the above - mentioned method 1400, the heat generated during a single - phase fault can be spread across all the FETs in the inverter circuit 1300 to ensure safety and reduce damage when a fault occurs. Additionally, any one or more of the steps 1410 - 1450 performed in method 1400 can occur within certain time ranges between the steps, such as nanoseconds, milliseconds, or any other time value. In some embodiments, the control circuit in the system can provide control signals to selectively open or close the switches (e.g., FETs) in the inverter circuit 1300. The control circuit can also provide squib signals to activate one or both of the squib fuses F1 and F2. In some other embodiments, the system can further include a plurality of control circuits to send squib signals for activating the squib fuses F1 and F2 and control signals for controlling one or more FETs in the inverter circuit 1300. It should be understood that any of the steps discussed herein do not necessarily have to be performed in a specific number of stages or within a certain stage. The steps can be performed throughout method 1400 as needed. Additionally, the above description includes exemplary steps and / or operations, but operations can be appropriately added, replaced, reordered, and / or eliminated without departing from the spirit and scope of the present disclosure. It should be understood that the method for controlling the inverter circuit and the inverter circuit disclosed in various embodiments can also be used in various fields or systems, including but not limited to the fields of automotive, hybrid and electric vehicles, electric motors, etc.

[0141] Various embodiments herein are described in the general context of method steps or processes. In one aspect, these method steps or processes may be implemented by an integrated circuit that includes circuitry for performing a method of controlling an inverter circuit. The circuitry may be configured to perform the steps or processes of the above-described method. For example, the circuitry may include one or more controllers, one or more processors, or a combination thereof to control the inverter circuit disclosed in various embodiments of the present disclosure.

[0142] Various embodiments herein are described in the general context of method steps or processes. In one aspect, these method steps or processes are implemented by a computer program product embodied in a transient or non-transient computer-readable medium that stores computer-executable instructions, such as program code, executed by one or more processors or one or more controllers in a system. The computer-readable medium may include removable storage devices and non-removable storage devices, including but not limited to read-only memory devices (ROM), random access memory devices (RAM), optical discs (CD), digital versatile discs (DVD), and the like.

[0143] Generally speaking, program modules may include routines, programs, objects, components, data structures, etc. that perform specific tasks or implement specific abstract data types. The computer-executable instructions, associated data structures, and program modules represent examples of program code for performing the steps of the methods disclosed herein. Specific sequences of such executable instructions or associated data structures represent examples of corresponding behaviors for implementing the functions described in such steps or processes.

[0144] Reference Figure 15 , which shows a graph of the bus voltage of the HV DC bus versus time during a discharge cycle consistent with some embodiments of the present disclosure. As shown by curve 1510, during a conventional discharge process, the voltage, current, and charge can decay exponentially during capacitor discharge. On the other hand, as shown by curve 1520, in some embodiments of the present disclosure, a fast discharge can be achieved. Thus, at the beginning, the bus voltage Vbus can decay exponentially along the exponential decay curve, but once the bus voltage Vbus drops to a certain threshold V at time T T , the bus voltage Vbus can rapidly drop to zero instead of continuing along the exponential decay curve. As explained in the above embodiments, after the inverter circuit 1300 detects that the bus voltage Vbus is below a specific threshold V at time T T , various circuitry (e.g., Figure 13B the fast discharge component 1342 therein) can be used to rapidly discharge the remaining voltage.

[0145] Reference Figures 16A - 16D, these figures are graphs showing EMI noise consistent with some embodiments of the present disclosure. Figures 16A - 16B are graphs respectively showing the EMI noise detected on the positive and negative sides of a dual-inverter system consistent with some embodiments of the present disclosure. Figures 16C - 16D are graphs respectively showing the EMI noise detected on the positive and negative sides of a single-inverter system consistent with some embodiments of the present disclosure. As Figures 16A - 16D shown, in various embodiments disclosed in the present disclosure, the EMI noise in a single-inverter system is about 30 - 40 dB higher than that in the proposed system using two inverters with PWM signals staggered by substantially 180 degrees. Therefore, through the common-mode noise suppression / elimination achieved by a power propulsion system having two inverters, the volume and mass of the required electromagnetic compatibility (EMC) components can be reduced. As a result, a lightweight engine design can be achieved.

[0146] In the above specification, embodiments have been described with reference to many specific details, which may vary according to the implementation. Certain adjustments and modifications can be made to the described embodiments. The order of steps shown in the figures is also considered for illustrative purposes only and is not intended to be limited to any specific order of steps. Therefore, those skilled in the art can understand that when implementing the same method, these steps can be executed in a different order.

[0147] As used herein, unless otherwise explicitly stated, the term "or" encompasses all possible combinations, unless infeasible. For example, if it is stated that a module can include A or B, then unless otherwise specifically stated or infeasible, the module can include A, or B, or A and B. As a second example, if it is stated that a module can include A, B, or C, then unless otherwise specifically stated or infeasible, the module can include A, or B, or C, or A and B, or A and C, or B and C, or A and B and C.

[0148] In the drawings and the specification, exemplary embodiments have been disclosed. It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed devices, systems, and related methods. Considering the specification and practice of the disclosed devices, systems, and related methods, other embodiments will be apparent to those skilled in the art. The specification and examples are intended to be merely exemplary, and the true scope is indicated by the appended claims and their equivalents.

[0149] The embodiments can be further described using the following terms:

[0150] Clause set 1

[0151] 1. A propulsion system for an aircraft, the propulsion system comprising:

[0152] An electric motor configured to drive one or more propellers of the aircraft; and

[0153] A capacitor configured to stabilize a direct current (DC) bus voltage;

[0154] A first inverter circuit coupled to the capacitor and configured to convert the DC bus voltage on a first bus of the first inverter circuit into an alternating current (AC) voltage based on a first pulse width modulation (PWM) vector to drive a first set of stator windings of the electric motor; and

[0155] A second inverter circuit coupled to the capacitor and configured to convert the DC bus voltage on a second bus of the second inverter circuit into an AC voltage based on a second PWM vector to drive a second set of stator windings of the electric motor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors.

[0156] 2. The propulsion system according to clause 1, wherein the first set of stator windings and the second set of stator windings are offset by substantially 180 degrees.

[0157] 3. The propulsion system according to clause 1 or 2, wherein a centered space vector modulation is used to control the first inverter circuit.

[0158] 4. The propulsion system according to any one of the preceding clauses, wherein an inverted centered space vector modulation is used to control the second inverter circuit.

[0159] 5. The propulsion system according to any one of the preceding clauses, wherein the first inverter circuit is configured to output a first set of three-phase AC voltages, and the second inverter circuit is configured to output a second set of three-phase AC voltages.

[0160] 6. The propulsion system according to clause 5, wherein the phases of the first set of three-phase AC voltages and the corresponding phases of the second set of three-phase AC voltages are two interleaved phases having a phase shift of substantially 180 degrees.

[0161] 7. The propulsion system according to any one of the preceding clauses, the propulsion system further comprising:

[0162] A DC common mode filter coupled to the capacitor and configured to reduce common mode signals on the DC sides of the first inverter circuit and the second inverter circuit.

[0163] 8. The propulsion system according to any one of the preceding clauses, the propulsion system further comprising:

[0164] One or more AC common mode chokes, the one or more AC common mode chokes being coupled to the AC side of the first inverter circuit or the second inverter circuit to reduce common mode signals.

[0165] Clause set 2

[0166] 9. A method for controlling a propulsion system of an aircraft, the method comprising:

[0167] Stabilizing a direct current (DC) bus voltage through a capacitor;

[0168] Converting the DC bus voltage to an alternating current (AC) voltage by a first inverter circuit coupled to the capacitor according to a first pulse width modulation (PWM) vector to drive a first set of stator windings of an electric motor;

[0169] Converting the DC bus voltage to an AC voltage by a second inverter circuit coupled to the capacitor in response to a second PWM vector to drive a second set of stator windings of the electric motor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors; and

[0170] Driving one or more propellers of the aircraft by the electric motor.

[0171] 10. The method according to clause 9, wherein the first set of stator windings and the second set of stator windings are offset by substantially 180 degrees.

[0172] 11. The method according to clause 9 or 10, the method further comprising:

[0173] Using centered space vector modulation to control the first inverter circuit.

[0174] 12. The method according to any one of clauses 9 to 11, the method further comprising:

[0175] Using inverted centered space vector modulation to control the second inverter circuit.

[0176] 13. The method according to any one of clauses 9 to 12, the method further comprising:

[0177] Outputting a first set of three-phase AC voltages through the first inverter circuit to drive the first set of stator windings; and

[0178] Outputting a second set of three-phase AC voltages through the second inverter circuit to drive the second set of stator windings.

[0179] 14. The method according to clause 13, wherein the phases of the first set of three-phase AC voltages and the corresponding phases of the second set of three-phase AC voltages are two interleaved phases having a phase shift of substantially 180 degrees.

[0180] 15. The method according to any one of clauses 9 to 14, the method further comprising:

[0181] Reducing the common-mode signal on the DC side of the first inverter circuit and the second inverter circuit by a DC common-mode filter coupled to the capacitor.

[0182] 16. The method according to any one of clauses 9 to 15, the method further comprising:

[0183] Reducing the common-mode signal on the AC side of the first inverter circuit and the second inverter circuit by one or more AC common-mode chokes coupled to the AC side of the first inverter circuit or the second inverter circuit.

[0184] Clause set 3

[0185] 17. An integrated circuit, the integrated circuit comprising circuitry for performing a method for controlling a propulsion system of an aircraft, the circuitry configured to:

[0186] Control a first inverter circuit coupled to a capacitor to convert a DC bus voltage into an alternating current (AC) voltage according to a first pulse width modulation (PWM) vector to drive a first set of stator windings of an electric motor; and

[0187] Control a second inverter circuit coupled to the capacitor to convert the DC bus voltage into an AC voltage in response to a second PWM vector to drive a second set of stator windings of the electric motor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors to drive one or more propellers of the aircraft through the electric motor.

[0188] 18. The integrated circuit according to clause 17, wherein the first set of stator windings and the second set of stator windings are offset by substantially 180 degrees.

[0189] 19. The integrated circuit according to clause 17 or 18, wherein the circuitry is further configured to control the first inverter circuit using centered space vector modulation.

[0190] 20. The integrated circuit according to any one of clauses 17 to 19, wherein the circuitry is further configured to control the second inverter circuit using inverted centered space vector modulation.

[0191] 21. The integrated circuit according to any one of clauses 17 to 20, wherein the circuit system is further configured to:

[0192] control the first inverter circuit to output a first set of three-phase AC voltages through the first inverter circuit to drive the first set of stator windings; and

[0193] control the second inverter circuit to output a second set of three-phase AC voltages through the second inverter circuit to drive the second set of stator windings.

[0194] 22. The integrated circuit according to clause 21, wherein the phases of the first set of three-phase AC voltages and the corresponding phases of the second set of three-phase AC voltages are two interleaved phases having a phase shift of substantially 180 degrees.

[0195] 23. The integrated circuit according to any one of clauses 17 to 22, wherein a DC common-mode filter is coupled to the capacitor and is configured to reduce the common-mode signals on the DC sides of the first inverter circuit and the second inverter circuit.

[0196] 24. The integrated circuit according to any one of clauses 17 to 23, wherein one or more AC common-mode chokes are coupled to the AC sides of the first inverter circuit or the second inverter circuit to reduce common-mode signals.

[0197] Clause set 4

[0198] 25. An inverter circuit for a propulsion system of an aircraft, the inverter circuit comprising:

[0199] a capacitor configured to stabilize a direct current (DC) bus voltage;

[0200] a plurality of switches forming a plurality of phase bridge arms, wherein at least one of the phase bridge arms includes an upper switch arranged between the positive terminal of the capacitor and the AC output terminal of the phase bridge arm and a lower switch arranged between the negative terminal of the capacitor and the AC output terminal of the phase bridge arm; and

[0201] a first discharge circuit coupled in parallel to the capacitor and configured to provide a first discharge path for discharging the energy stored in the capacitor,

[0202] wherein in response to the DC bus voltage being lower than a first threshold in a fault condition associated with the inverter circuit, the plurality of switches are controlled to short-circuit the capacitor.

[0203] 26. The inverter circuit according to clause 25, wherein the first discharge circuit includes a discharge resistor and a first switch connected in series to the discharge resistor, and in response to a single-phase short circuit, the inverter circuit is configured to discharge the capacitor by closing the first switch.

[0204] 27. The inverter circuit according to clause 25 or 26, the inverter circuit further comprising:

[0205] A second discharge circuit, the second discharge circuit being coupled in parallel to the capacitor and configured to provide a second discharge path for discharging the energy stored in the capacitor.

[0206] 28. The inverter circuit according to clause 27, wherein the second discharge circuit includes a discharge component and a second switch connected in series.

[0207] 29. The inverter circuit according to clause 28, wherein the discharge component includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance value device, or any combination thereof.

[0208] 30. The inverter circuit according to any one of clauses 25 to 29, the inverter circuit further comprising:

[0209] A comparator circuit, the comparator circuit being configured to monitor the DC bus voltage across the capacitor to determine whether to discharge the energy stored in the capacitor through the second discharge circuit.

[0210] 31. The inverter circuit according to clause 30, wherein the comparator circuit includes:

[0211] A resistive voltage divider, the resistive voltage divider including resistors connected in series to provide a first voltage, wherein the first voltage is a fixed fraction of the DC bus voltage; and

[0212] A comparator, the comparator being coupled to the resistive voltage divider and configured to:

[0213] Compare the first voltage output from the resistive voltage divider with a reference voltage, and

[0214] Output an output signal in response to the first voltage being lower than the reference voltage.

[0215] 32. The inverter circuit according to clause 31, wherein the comparator circuit further includes:

[0216] A logic circuit, the logic circuit being coupled to the comparator and configured to:

[0217] Receiving the output signal from the comparator and a command signal from a control circuit, and

[0218] Outputting a control signal according to the output signal and the command signal to selectively turn on the second switch of the second discharge circuit.

[0219] Clause set 5

[0220] 33. A method for controlling an inverter circuit, the method comprising:

[0221] Detecting whether a fault occurs on one of a plurality of switches in the inverter circuit;

[0222] In response to detecting a single-phase short circuit fault, disconnecting the inverter circuit from a power source;

[0223] After disconnecting the inverter circuit from the power source, providing a first discharge path through a first discharge circuit to discharge the DC bus voltage across the capacitor of the inverter circuit; and

[0224] In response to the DC bus voltage being lower than a first threshold, controlling the plurality of switches in the inverter circuit to short-circuit the capacitor.

[0225] 34. The method according to clause 33, wherein the first discharge circuit includes a discharge resistor and a first switch connected in series to the discharge resistor, the method comprising: discharging the capacitor by closing the first switch via the inverter circuit in response to a single-phase short circuit.

[0226] 35. The method according to clause 33 or 34, the method further comprising:

[0227] Providing a second discharge path through a second discharge circuit coupled in parallel to the capacitor to discharge the DC bus voltage.

[0228] 36. The method according to clause 35, wherein the second discharge circuit includes a discharge component and a second switch connected in series.

[0229] 37. The method according to clause 36, wherein the discharge component includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance value device, or any combination thereof.

[0230] 38. The method according to any one of clauses 33 to 37, the method further comprising:

[0231] The DC bus voltage across the capacitor is monitored by a comparator circuit to determine whether to discharge the energy stored in the capacitor through the second discharge circuit.

[0232] 39. The method according to any one of clauses 33 to 38, the method further comprising:

[0233] providing a first voltage through a resistive voltage divider including resistors connected in series, wherein the first voltage is a fixed fraction of the bus voltage; and

[0234] comparing the first voltage with a reference voltage by a comparator to output an output signal in response to the first voltage being lower than the reference voltage.

[0235] 40. The method according to clause 39, the method further comprising:

[0236] receiving the output signal from the comparator and a command signal from a control circuit through a logic circuit; and

[0237] outputting a control signal through the logic circuit to selectively turn on the second switch of the second discharge circuit according to the output signal and the command signal.

[0238] Clause set 6

[0239] 41. An integrated circuit, the integrated circuit includes circuitry for performing a method for controlling an inverter circuit, the circuitry being configured to perform a method comprising:

[0240] detecting whether a fault occurs on one of a plurality of switches in the inverter circuit;

[0241] disconnecting the inverter circuit from a power source in response to detecting a single-phase short circuit fault;

[0242] after disconnecting the inverter circuit from the power source, controlling the inverter circuit to provide a first discharge path through a first discharge circuit to discharge the DC bus voltage across the capacitor of the inverter circuit; and

[0243] in response to the DC bus voltage being lower than a first threshold, controlling the inverter circuit to control the plurality of switches in the inverter circuit to short-circuit the capacitor.

[0244] 42. The integrated circuit according to clause 41, wherein the first discharge circuit includes a discharge resistor and a first switch connected in series to the discharge resistor, and in response to a single-phase short circuit, the circuit system is configured to control the inverter circuit to discharge the capacitor by closing the first switch.

[0245] 43. The integrated circuit according to clause 41 or 42, wherein the circuit system is configured to perform:

[0246] Controlling the inverter circuit to provide a second discharge path through a second discharge circuit coupled in parallel to the capacitor to discharge the DC bus voltage.

[0247] 44. The integrated circuit according to clause 43, wherein the second discharge circuit includes a discharge component and a second switch connected in series.

[0248] 45. The integrated circuit according to clause 44, wherein the discharge component includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance value device, or any combination thereof.

[0249] 46. The integrated circuit according to any one of clauses 43 to 45, wherein the DC bus voltage across the capacitor is monitored by a comparator circuit to determine whether to discharge the energy stored in the capacitor through the second discharge circuit.

[0250] 47. The integrated circuit according to any one of clauses 41 to 46, wherein the DC bus voltage across the capacitor is monitored by:

[0251] Providing a first voltage through a resistive voltage divider including a resistor connected in series, wherein the first voltage is a fixed fraction of the bus voltage; and

[0252] Comparing the first voltage with a reference voltage by a comparator to output an output signal in response to the first voltage being lower than the reference voltage.

[0253] 48. The integrated circuit according to clause 47, wherein the DC bus voltage across the capacitor is monitored by:

[0254] Receiving the output signal from the comparator and a command signal from a control circuit by a logic circuit; and

[0255] Outputting a control signal by the logic circuit to selectively turn on the second switch of the second discharge circuit according to the output signal and the command signal.

[0256] Clause set 7

[0257] 49. An inverter circuit for a propulsion system of an aircraft, the inverter circuit comprising:

[0258] A capacitor configured to stabilize a direct current (DC) bus voltage;

[0259] A plurality of switches forming a plurality of phase bridge arms, wherein at least one of the phase bridge arms includes an upper switch disposed between a positive terminal of the capacitor and an AC output terminal of the phase bridge arm and a lower switch disposed between a negative terminal of the capacitor and the AC output terminal of the phase bridge arm;

[0260] A first discharge circuit coupled in parallel to the capacitor and configured to provide a first discharge path for discharging energy stored in the capacitor; and

[0261] A second discharge circuit coupled in parallel to the capacitor and configured to provide a second discharge path for discharging energy stored in the capacitor in response to the DC bus voltage being below a threshold in a fault condition associated with the inverter circuit.

[0262] 50. The inverter circuit according to clause 49, wherein the first discharge circuit includes a discharge resistor and a first switch connected in series to the discharge resistor, and in response to a single-phase short circuit, the inverter circuit is configured to discharge the capacitor by closing the first switch.

[0263] 51. The inverter circuit according to clause 49 or 50, wherein the second discharge circuit includes a discharge component and a second switch connected in series.

[0264] 52. The inverter circuit according to clause 51, wherein the discharge component includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance value device, or any combination thereof.

[0265] 53. The inverter circuit according to any one of clauses 49 to 52, the inverter circuit further comprising:

[0266] A comparator circuit configured to monitor the DC bus voltage across the capacitor to determine whether to discharge the energy stored in the capacitor through the second discharge circuit.

[0267] 54. The inverter circuit according to clause 53, wherein the comparator circuit includes:

[0268] A resistive voltage divider, the resistive voltage divider including resistors connected in series to provide a first voltage, where the first voltage is a fixed fraction of the DC bus voltage; and

[0269] A comparator, the comparator being coupled to the resistive voltage divider and configured to:

[0270] Compare the first voltage output from the resistive voltage divider with a reference voltage, and

[0271] Output an output signal in response to the first voltage being lower than the reference voltage.

[0272] 55. The inverter circuit according to clause 54, wherein the comparator circuit further includes:

[0273] A logic circuit, the logic circuit being coupled to the comparator and configured to:

[0274] Receive the output signal from the comparator and a command signal from a control circuit, and

[0275] Output a control signal according to the output signal and the command signal to selectively turn on the second switch of the second discharge circuit.

[0276] Clause set 8

[0277] 56. A method for controlling an inverter circuit, the method including:

[0278] Detect whether a fault occurs on one of a plurality of switches in the inverter circuit;

[0279] In response to detecting a single-phase short circuit fault, disconnect the inverter circuit from the power source; and

[0280] After disconnecting the inverter circuit from the power source, discharge the bus voltage across the capacitor of the inverter circuit by:

[0281] In response to confirming that the inverter circuit is disconnected from the power source, use a first discharge circuit to provide a first discharge path; and

[0282] In response to the bus voltage being lower than a threshold, use a second discharge circuit to provide a second discharge path in parallel with the first discharge path.

[0283] 57. The method according to clause 56, wherein the first discharge circuit includes a discharge resistor and a first switch connected in series to the discharge resistor, and discharging the bus voltage across the capacitor further includes:

[0284] In response to confirming that the inverter circuit is disconnected from the power source, close the first switch in the first discharge circuit to provide the first discharge path.

[0285] 58. The method according to clause 56 or 57, wherein the second discharge circuit includes a discharge component and a second switch connected in series, and discharging the bus voltage across the capacitor further includes:

[0286] In response to the bus voltage being lower than the threshold, close the second switch in the second discharge circuit to provide the second discharge path in parallel with the first discharge path.

[0287] 59. The method according to clause 58, wherein the discharge component includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance value device, or any combination thereof.

[0288] 60. The method according to any one of clauses 56 to 59, wherein discharging the bus voltage across the capacitor further includes:

[0289] Monitoring the bus voltage across the capacitor through a comparator circuit to determine whether to discharge the energy stored in the capacitor through the second discharge circuit.

[0290] 61. The method according to any one of clauses 56 to 60, wherein discharging the bus voltage across the capacitor further includes:

[0291] Providing a first voltage through a resistive voltage divider including a resistor connected in series, wherein the first voltage is a fixed fraction of the bus voltage; and

[0292] Comparing the first voltage with a reference voltage through a comparator to output an output signal in response to the first voltage being lower than the reference voltage.

[0293] 62. The method according to clause 61, wherein discharging the bus voltage across the capacitor further includes:

[0294] Receiving the output signal from the comparator and a command signal from a control circuit through a logic circuit; and

[0295] According to the output signal and the command signal, output a control signal through the logic circuit to selectively turn on the second switch of the second discharge circuit.

[0296] Clause set 9

[0297] 63. An integrated circuit, the integrated circuit including circuitry for performing a method for controlling an inverter circuit, the circuitry configured to:

[0298] Detect whether a fault has occurred on one of a plurality of switches in the inverter circuit;

[0299] In response to detecting a single-phase short circuit fault, disconnect the inverter circuit from a power source; and

[0300] After disconnecting the inverter circuit from the power source, control the inverter circuit to discharge the bus voltage across the capacitor of the inverter circuit by:

[0301] In response to confirming that the inverter circuit is disconnected from the power source, use a first discharge circuit to provide a first discharge path; and

[0302] In response to the bus voltage being lower than a threshold, use a second discharge circuit to provide a second discharge path in parallel with the first discharge path.

[0303] 64. The integrated circuit according to clause 63, wherein the first discharge circuit includes a discharge resistor and a first switch connected in series to the discharge resistor, and the inverter circuit is controlled to discharge the bus voltage across the capacitor by:

[0304] In response to confirming that the inverter circuit is disconnected from the power source, close the first switch in the first discharge circuit to provide the first discharge path.

[0305] 65. The integrated circuit according to clause 63 or 64, wherein the second discharge circuit includes a discharge component and a second switch connected in series, and the inverter circuit is controlled to discharge the bus voltage across the capacitor by:

[0306] In response to the bus voltage being lower than the threshold, close the second switch in the second discharge circuit to provide the second discharge path.

[0307] 66. The method according to clause 65, wherein the discharge component includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance value device, or any combination thereof.

[0308] 67. The integrated circuit according to any one of clauses 63 to 66, wherein the bus voltage across the capacitor is discharged by:

[0309] The bus voltage across the capacitor is monitored by a comparator circuit to determine whether to discharge the energy stored in the capacitor through the second discharge circuit.

[0310] 68. The integrated circuit according to any one of clauses 63 to 67, wherein the bus voltage across the capacitor is discharged by:

[0311] providing a first voltage through a resistive voltage divider including resistors connected in series, wherein the first voltage is a fixed fraction of the bus voltage; and

[0312] comparing the first voltage with a reference voltage by a comparator to output an output signal in response to the first voltage being lower than the reference voltage.

[0313] 69. The integrated circuit according to any one of clauses 63 to 68, wherein the bus voltage across the capacitor is discharged by:

[0314] receiving the output signal from the comparator and a command signal from a control circuit by a logic circuit; and

[0315] outputting, by the logic circuit, a control signal to selectively turn on the second switch of the second discharge circuit according to the output signal and the command signal.

[0316] The embodiments disclosed herein are intended to be non-limiting. Those of ordinary skill in the art will understand that certain components and the configuration of components can be modified without departing from the scope of the disclosed embodiments.

Claims

1. A propulsion system for an aircraft, the propulsion system comprising: an electric motor configured to drive one or more propellers of the aircraft; a capacitor configured to stabilize a direct current (DC) bus voltage; a first inverter circuit coupled to the capacitor and configured to convert the DC bus voltage on a first bus of the first inverter circuit into an alternating current (AC) voltage based on a first pulse width modulation (PWM) vector to drive a first set of stator windings of the electric motor; and a second inverter circuit coupled to the capacitor and configured to convert the DC bus voltage on a second bus of the second inverter circuit into an AC voltage based on a second PWM vector to drive a second set of stator windings of the electric motor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors.

2. The propulsion system according to claim 1, wherein the first set of stator windings and the second set of stator windings are offset by substantially 180 degrees.

3. The propulsion system according to claim 1 or 2, wherein a centered space vector modulation is used to control the first inverter circuit.

4. The propulsion system according to any one of claims 1 to 3, wherein an inverted centered space vector modulation is used to control the second inverter circuit.

5. The propulsion system according to any one of claims 1 to 4, wherein the first inverter circuit is configured to output a first set of three-phase AC voltages, and the second inverter circuit is configured to output a second set of three-phase AC voltages.

6. The propulsion system according to claim 5, wherein a phase of the first set of three-phase AC voltages and a corresponding phase of the second set of three-phase AC voltages are two interleaved phases having a phase shift of substantially 180 degrees.

7. The propulsion system according to any one of claims 1 to 6, the propulsion system further comprising: a DC common mode filter coupled to the capacitor and configured to reduce common mode signals on a DC side of the first inverter circuit and the second inverter circuit.

8. The propulsion system according to any one of claims 1 to 7, the propulsion system further comprising: one or more AC common mode chokes coupled to an AC side of the first inverter circuit or the second inverter circuit to reduce common mode signals.

9. A method for controlling a propulsion system of an aircraft, the method comprising: stabilizing a direct current (DC) bus voltage by a capacitor; converting the DC bus voltage into an alternating current (AC) voltage by a first inverter circuit coupled to the capacitor based on a first pulse width modulation (PWM) vector to drive a first set of stator windings of an electric motor; In response to a second PWM vector, convert the DC bus voltage to an AC voltage by a second inverter circuit coupled to the capacitor to drive a second set of stator windings of the electric motor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors; And Drive one or more propellers of the aircraft by the electric motor.

10. The method according to claim 9, wherein the first set of stator windings and the second set of stator windings are offset by substantially 180 degrees.

11. The method according to claim 9 or 10, the method further comprising: Using centered space vector modulation to control the first inverter circuit.

12. The method according to any one of claims 9 to 11, the method further comprising: Using inverted centered space vector modulation to control the second inverter circuit.

13. The method according to any one of claims 9 to 12, the method further comprising: Output a first set of three-phase AC voltages through the first inverter circuit to drive the first set of stator windings; And Output a second set of three-phase AC voltages through the second inverter circuit to drive the second set of stator windings.

14. The method according to claim 13, wherein the phases of the first set of three-phase AC voltages and the corresponding phases of the second set of three-phase AC voltages are two interleaved phases having a phase shift of substantially 180 degrees.

15. The method according to any one of claims 9 to 14, the method further comprising: Reduce the common-mode signals on the DC sides of the first inverter circuit and the second inverter circuit by a DC common-mode filter coupled to the capacitor.

16. The method according to any one of claims 9 to 15, the method further comprising: Reduce the common-mode signals on the AC sides of the first inverter circuit and the second inverter circuit by one or more AC common-mode chokes coupled to the AC side of the first inverter circuit or the second inverter circuit.

17. An integrated circuit, the integrated circuit comprising circuitry for performing a method for controlling a propulsion system of an aircraft, the circuitry configured to: Control a first inverter circuit coupled to a capacitor to convert a DC bus voltage to an alternating current (AC) voltage according to a first pulse width modulation (PWM) vector to drive a first set of stator windings of an electric motor; and Control a second inverter circuit coupled to the capacitor to convert the DC bus voltage to an AC voltage in response to a second PWM vector to drive a second set of stator windings of the electric motor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors, to drive one or more propellers of the aircraft by the electric motor.

18. The integrated circuit according to claim 17, wherein the first set of stator windings and the second set of stator windings are offset by substantially 180 degrees.

19. The integrated circuit according to claim 17 or 18, wherein the circuit system is further configured to control the first inverter circuit using center - aligned space vector modulation.

20. The integrated circuit according to any one of claims 17 to 19, wherein the circuit system is further configured to control the second inverter circuit using inverted center - aligned space vector modulation.

21. The integrated circuit according to any one of claims 17 to 20, wherein the circuit system is further configured to: control the first inverter circuit to drive the first set of stator windings by outputting a first set of three - phase AC voltages through the first inverter circuit; and control the second inverter circuit to drive the second set of stator windings by outputting a second set of three - phase AC voltages through the second inverter circuit.

22. The integrated circuit according to claim 21, wherein the phase of the first set of three - phase AC voltages and the corresponding phase of the second set of three - phase AC voltages are two interleaved phases having a phase shift of substantially 180 degrees.

23. The integrated circuit according to any one of claims 17 to 22, wherein a DC common - mode filter is coupled to the capacitor and is configured to reduce the common - mode signal on the DC side of the first inverter circuit and the second inverter circuit.

24. The integrated circuit according to any one of claims 17 to 23, wherein one or more AC common - mode chokes are coupled to the AC side of the first inverter circuit or the second inverter circuit to reduce the common - mode signal.

25. An inverter circuit for a propulsion system of an aircraft, the inverter circuit comprising: a capacitor configured to stabilize a direct - current (DC) bus voltage; a plurality of switches forming a plurality of phase bridge arms, wherein at least one of the phase bridge arms includes an upper switch disposed between the positive terminal of the capacitor and the AC output terminal of the phase bridge arm and a lower switch disposed between the negative terminal of the capacitor and the AC output terminal of the phase bridge arm; and a first discharge circuit coupled in parallel to the capacitor and configured to provide a first discharge path for discharging the energy stored in the capacitor, wherein in response to the DC bus voltage being lower than a first threshold in a fault condition associated with the inverter circuit, the plurality of switches are controlled to short - circuit the capacitor.

26. The inverter circuit according to claim 25, wherein the first discharge circuit includes a discharge resistor and a first switch connected in series to the discharge resistor, and in response to a single - phase short - circuit, the inverter circuit is configured to discharge the capacitor by closing the first switch.

27. The inverter circuit according to claim 25 or 26, the inverter circuit further comprising: a second discharge circuit coupled in parallel to the capacitor and configured to provide a second discharge path for discharging the energy stored in the capacitor.

28. The inverter circuit according to claim 27, wherein the second discharge circuit includes a discharge component and a second switch connected in series.

29. The inverter circuit according to claim 28, wherein the discharge component includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance device, or any combination thereof.

30. The inverter circuit according to any one of claims 25 to 29, the inverter circuit further comprising: a comparator circuit configured to monitor the DC bus voltage across the capacitor to determine whether to discharge the energy stored in the capacitor through the second discharge circuit.

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