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 to drive a first set of stator windings of the electric motor in response to a first pulse-width modulation (PWM) vector, 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, where the first PWM vector and the second PWM vector are substantially equal and opposite vectors.

Description

[Technical Field]

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS This disclosure claims priority to U.S. patent application Ser. No. 18 / 363,535, entitled "INVERTER CIRCUITS AND ELECTRICAL PROPULSION SYSTEMS FOR EVTOL AIRCRAFT," filed August 1, 2023, which in turn claims priority to U.S. Provisional Application Ser. No. 63 / 378,536, entitled "Tilt Rotor Systems and Methods for eVTOL Aircraft," filed October 6, 2022, and U.S. Provisional Application Ser. No. 63 / 378,680, entitled "Systems and Methods for Improved Propulsion Systems for eVTOL Aircraft," filed October 7, 2022, the contents of which are incorporated herein by reference in their entirety for all purposes.

[0002] The present disclosure relates generally to the field of powered air vehicles. More specifically, but not exclusively, the present disclosure relates to innovations in tiltrotor aircraft using electric propulsion systems. Certain aspects of the present disclosure generally relate to improvements in electric propulsion systems for tiltrotor aircraft. Other aspects of the present disclosure generally relate to improvements in power inverters that may be used in other types of vehicles but provide particular advantages in air vehicles. Summary of the Invention

[0003] An embodiment of the present disclosure provides 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.

[0004] An embodiment of the present disclosure provides a method for controlling a propulsion system for an aircraft. The method may include stabilizing a direct current (DC) bus voltage with a capacitor; converting the DC bus voltage to an alternating current (AC) voltage according to a first pulse-width modulation (PWM) vector with a first inverter circuit coupled to the capacitor to drive a first set of stator windings of an electric motor; converting the DC bus voltage to an AC voltage with a second inverter circuit coupled to the capacitor to drive a second set of stator windings of the electric motor in response to a second PWM vector, where the first PWM vector and the second PWM vector are substantially equal and opposite vectors; and driving one or more propellers of the aircraft with the electric motor. Other embodiments may include corresponding integrated circuits, computer systems, apparatus, and computer programs recorded on one or more computer storage devices, each configured to perform the operations of the method.

[0005] An embodiment of the present disclosure provides an inverter circuit. The inverter circuit may include a capacitor configured to stabilize a direct current (DC) bus voltage and a plurality of switches forming a plurality of phase legs, 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, 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. The plurality of switches are controlled to short-circuit the capacitor in response to the DC bus voltage being lower than a threshold value in a fault condition.

[0006] An embodiment of the present disclosure provides a method for controlling an inverter circuit. The method may include detecting whether a fault occurs in one of a plurality of switches in the inverter circuit, disconnecting the inverter circuit from a power source in response to detecting a single-phase short-circuit fault, providing, with a first discharge circuit, a first discharge path for discharging a DC bus voltage across a 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 lower than a first threshold. Other embodiments may include corresponding integrated circuits, computer systems, apparatus, and computer programs recorded on one or more computer storage devices, each configured to perform the operations of the method.

[0007] An embodiment of the present disclosure provides an inverter circuit that may include a capacitor configured to stabilize a direct current (DC) bus voltage and a plurality of switches forming a plurality of phase legs, 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, 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 discharging energy stored in the capacitor in response to the DC bus voltage being lower than a threshold value in a fault condition.

[0008] An embodiment of the present disclosure provides a method for controlling an inverter circuit. The method may include detecting whether a fault occurs in one of a plurality of switches in the inverter circuit; disconnecting the inverter circuit from a power source in response to detecting a single-phase short-circuit fault; and, after the inverter circuit is disconnected from the power source, discharging a bus voltage across a capacitor of the inverter circuit by providing a first discharge path using a first discharge circuit in response to determining that the inverter circuit is disconnected from the power source, and by providing a second discharge path in parallel to the first discharge path using a second discharge circuit in response to the bus voltage being lower than a threshold. Other embodiments may include corresponding integrated circuits, computer systems, apparatus, and computer programs recorded on one or more computer storage devices, each configured to perform the operations of the method. [Brief explanation of the drawings]

[0009] [Figure 1] 1 is an illustration of a perspective view of an exemplary VTOL aircraft consistent with certain embodiments of the present disclosure. [Figure 2] FIG. 1 is another illustration of a perspective view of an exemplary VTOL aircraft in an alternative configuration consistent with certain embodiments of the present disclosure. [Figure 3] 1 is an illustration of a top view of an exemplary VTOL aircraft, consistent with certain embodiments of the present disclosure. [Figure 4] FIG. 1 is a schematic diagram illustrating exemplary propeller rotation of a VTOL aircraft, consistent with certain embodiments of the present disclosure. [Figure 5] FIG. 1 is a schematic diagram illustrating exemplary power connections in a VTOL aircraft, consistent with certain embodiments of the present disclosure. [Figure 6] FIG. 1 is a block diagram illustrating an example architecture and design of an electric propulsion unit for a VTOL aircraft, consistent with an embodiment of the present disclosure. [Figure 7] FIG. 1 is a schematic diagram illustrating an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with an embodiment of the present disclosure. [Figure 8A] 1 is an illustration of an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with an embodiment of the present disclosure. [Figure 8B] 1 is an illustration of an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with an embodiment of the present disclosure. [Figure 8C] 1 is an illustration of an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with an embodiment of the present disclosure. [Figure 9] FIG. 1 is a schematic diagram illustrating an exemplary ascent electric propulsion system for a VTOL aircraft consistent with an embodiment of the present disclosure. [Figure 10A] 1 is an illustration of an exemplary electric ascent propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 10B] 1 is an illustration of an exemplary electric ascent propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 11] FIG. 1 illustrates a portion of an electric propulsion system for a vertical take-off and landing (VTOL) aircraft, consistent with some embodiments of the present disclosure. [Figure 12A] FIG. 12 illustrates PWM vectors for controlling inverter circuits in the electric propulsion system of FIG. 11 , consistent with some embodiments of the present disclosure. [Figure 12B]FIG. 2 is a diagram illustrating three-phase voltages output by an inverter circuit over one period, consistent with some embodiments of the present disclosure. [Figure 13A] FIG. 1 illustrates an inverter circuit for motor control in an electric propulsion system, consistent with some embodiments of the present disclosure. [Figure 13B] FIG. 10 illustrates another inverter circuit for motor control in an electric propulsion system, consistent with some embodiments of the present disclosure. [Figure 14] FIG. 13C illustrates an example flowchart of a method for controlling the inverter circuit of FIG. 13A or FIG. 13B, consistent with some embodiments of the present disclosure. [Figure 15] 1 is a graph illustrating bus voltage versus time for an HV DC bus during a discharge period, consistent with certain embodiments of the present disclosure. [Figure 16A] 1 is a graph illustrating EMI noise in a dual inverter system or a single inverter system, consistent with certain embodiments of the present disclosure. [Figure 16B] 1 is a graph illustrating EMI noise in a dual inverter system or a single inverter system, consistent with certain embodiments of the present disclosure. [Figure 16C] 1 is a graph illustrating EMI noise in a dual inverter system or a single inverter system, consistent with certain embodiments of the present disclosure. [Figure 16D] 1 is a graph illustrating EMI noise in a dual inverter system or a single inverter system, consistent with certain embodiments of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0010] The following disclosure provides different embodiments or examples for implementing different features of the provided subject matter. To simplify the disclosure, specific examples of components and arrangements are described below. These are, of course, merely examples and are not intended to be limiting. In addition, the disclosure may repeat reference numerals and / or letters in various examples. This repetition is for the purposes of simplicity and clarity and does not in itself dictate a relationship between the various embodiments and / or configurations discussed.

[0011] The terms used herein generally have their ordinary meanings in the art and in the specific context in which each term is used. The use of examples herein, including examples of any term discussed herein, is illustrative only and in no way limits the scope and meaning of the disclosure or any exemplary term. Likewise, the disclosure is not limited to the various embodiments provided herein.

[0012] Terms such as "first," "second," and the like may be used herein to describe various elements, but these elements should not be limited by these terms. These terms are used to distinguish one element from another. For example, a first element could be referred to as a second element, and similarly, a second element could be referred to as a first element without departing from the scope of the embodiment. As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items.

[0013] Additionally, spatially relative terms, such as "below," "lower than," "lower side," "above," "upper," and the like, may be used herein for ease of description to describe the relationship of one element or feature illustrated in the figures to another element(s) or feature(s). Spatially relative terms are intended to encompass different orientations of the device during use or operation in addition to the orientation shown in the figures. The device may be in other orientations (rotated 90 degrees or at other orientations) and the spatially relative descriptors used herein may be similarly interpreted accordingly.

[0014] As used herein, the term “coupled” may also be referred to as “electrically coupled,” and the term “connected” may also 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.

[0015] The present disclosure primarily addresses components of electric vertical take-off and landing (eVTOL) aircraft for use with non-traditional aircraft. For example, the eVTOL aircraft of the present disclosure may be intended for frequent (e.g., more than 50 flights per working day), short-duration flights (e.g., less than 100 miles per flight) over, into, and outside of populated areas. The aircraft may be intended to carry four to six passengers or commuters who expect a low-noise and low-vibration experience. Therefore, it may be desirable for the aircraft's components to be configured and designed to withstand frequent use without wear, for the components to generate less heat and vibration, and for the aircraft to include mechanisms for effectively controlling and managing the heat or vibration generated by the components. Furthermore, some of these aircraft may be intended to operate in close proximity to each other over congested metropolitan areas. Therefore, it may be desirable for the components to be configured and designed to generate low levels of noise both inside and outside the aircraft and to have various safety and backup mechanisms. For example, for safety reasons, it may be desirable for an aircraft to be propelled by a distributed propulsion system to avoid the risk of a single point of failure and to be capable of conventional takeoff and landing on a runway. Furthermore, it may be desirable for the aircraft to be able to safely take off and land vertically from relatively confined spaces (e.g., vertiports, parking lots, or driveways) while transporting approximately four to six passengers or commuters with associated baggage, as compared to traditional airport runways. These service requirements may impose design constraints on the size, weight, and operational efficiency (e.g., drag, energy use) of the aircraft, which may affect the design and configuration of aircraft components.

[0016] The disclosed embodiments provide new and improved configurations of aircraft components not observed in conventional aircraft and / or identified design criteria for components that differ from conventional aircraft components. Such alternative configurations and design criteria combine to address shortcomings and challenges of conventional components and result in the disclosed embodiments of various configurations and designs of eVTOL aircraft components.

[0017] In some embodiments, the disclosed eVTOL aircraft may be designed to be capable of both vertical and conventional takeoff and landing, with a distributed electric propulsion system that enables vertical flight, forward flight, and transition. Thrust may be generated by supplying high-voltage power to the electric engines of the distributed electric propulsion system, each of which may convert the high-voltage power into mechanical shaft power for rotating a propeller. The embodiments disclosed herein may involve optimizing the energy density of the electric propulsion system. The embodiments may include an electric engine connected to an onboard power source, which may include a device capable of storing energy, such as a battery or capacitor, or may include one or more systems for harnessing or generating electricity, such as a fuel-powered generator or a solar panel array. Some disclosed embodiments provide for component weight and space reduction in the aircraft, thereby increasing the aircraft's efficiency and performance. Focusing on safety in passenger transportation, the disclosed embodiments also implement new and improved safety protocols and system redundancies in the event of a failure to minimize any single point of failure in the aircraft's propulsion system. Some disclosed embodiments also provide new and improved approaches to meeting aviation and transportation laws and regulations. For example, the Federal Aviation Administration implements federal laws and regulations requiring safety components, such as fire barriers, adjacent to engines that use oil or other flammable materials in amounts above a threshold.

[0018] In a preferred embodiment, the distributed electric propulsion system may include 12 electric engines that may be mounted on forward and aft booms of the aircraft's wings. The forward electric engines may be tiltable in flight between a horizontally oriented position (e.g., to generate forward thrust) and a vertically oriented position (e.g., to generate vertical lift). The forward electric engines may be clockwise or counterclockwise with respect to the direction of propeller rotation. The aft electric engines may be fixed in a vertically oriented position (e.g., to generate vertical lift). The aft electric engines may also be clockwise or counterclockwise with respect to the direction of propeller rotation. In some embodiments, the aircraft may have various combinations of forward and aft electric engine configurations. For example, the aircraft may have six forward and six aft electric engines, four forward and four aft electric engines, or any other combination of forward and aft engines, including embodiments in which the number of forward and aft electric engines is unequal. In some embodiments, the aircraft may have four forward propellers and four aft propellers, at least four of which include tiltable propellers.

[0019] In a preferred embodiment, for vertical take-off and landing (VTOL) missions, the forward and aft electric engines can provide vertical thrust during take-off and landing. During flight phases when the aircraft is in forward flight mode, the forward electric engine can provide horizontal thrust, while the propeller of the aft electric engine can be retracted in a fixed position to minimize drag. The aft electric engine can be actively retracted with position monitoring. Transition from vertical to horizontal flight and vice versa can be achieved via a tilt propeller subsystem. The tilt propeller subsystem can redirect thrust between a primarily vertical direction during vertical flight mode and a generally horizontal direction during forward flight. A variable pitch mechanism can vary the collective angle of the propeller hub assembly blades of the forward electric engine for operation during hover, transition, and cruise phases.

[0020] In some embodiments, for conventional take-off and landing (CTOL) missions, the forward electric engine may provide horizontal thrust for fixed-wing take-off, cruise, and landing. In some embodiments, the aft electric engine may not be used to generate thrust during CTOL missions, and the aft propeller may be stowed in place.

[0021] In some embodiments, the electric engine may be housed in or connected to the boom of the aircraft and may include a motor, an inverter, and a gearbox. In some embodiments, the motor, inverter, and gearbox may be coupled together to share a central axis. In some embodiments, torque generated by the motor may be sent to the gearbox separately from the propeller of the propulsion system. In some embodiments, the gearbox may provide gear reduction and then send torque back to the propeller via a main shaft through a bearing located inside the motor. In some embodiments, the inverter may be attached to the back of the gearbox so that the main shaft does not pass through the inverter when outputting torque to the propeller. In some embodiments, the motor, gearbox, and inverter may be coupled together to maintain the motor, inverter, and / or gearbox while sharing a common heat exchanger using a coolant, such as oil. In some embodiments, the amount of oil used to lubricate and cool the electric engine may vary, including less than 1 quart, 2 quarts, 3 quarts, or any other metered amount of oil.

[0022] In some embodiments, the tilting propeller system may include a linear or rotary actuator for changing the orientation of the propulsion system during operation. In some embodiments, the pitch of the propulsion system may be changed as a function of 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 cooperate to provide a gear reduction that can orient the propulsion system. In some embodiments, the tilting propeller system may include a redundant configuration in which multiple motors, inverters, and gearboxes are present and cooperate 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 a motor, inverter, and gearbox of another portion of the configuration. In some embodiments, the gearbox configuration may also allow the tilting propeller system to maintain the orientation of the propulsion system with or without the assistance of additional power provided by the system.

[0023] In some embodiments, the electric propulsion systems described herein may generate thrust by supplying high-voltage (HV) electrical power to an electric engine, which converts the HV electrical power into mechanical shaft power used to rotate a propeller. As described above, the aircraft described herein may have multiple electric engines boom-mounted fore and aft of the wing. The amount of thrust generated by each electric engine may be controlled by torque commands from a flight control system (FCS) via a digital communication interface to each electric engine. Embodiments may include a forward electric engine, and may be capable of changing the orientation, or tilt, of the forward electric engine. Additional embodiments include a forward engine, which may be a clockwise (CW) or counterclockwise (CCW) type. The forward electric engine propulsion subsystem may consist of a multi-blade adjustable pitch propeller and a variable pitch subsystem.

[0024] In some embodiments, the aircraft may include an aft engine or lifter that may be of the clockwise (CW) or counterclockwise (CCW) type. Additional embodiments may include an aft electric engine that utilizes a multi-blade fixed pitch propeller.

[0025] As described herein, the orientation and use of the electric propulsion systems may vary throughout the operation of the aircraft. In some embodiments, during vertical takeoff and landing, the forward propulsion system and the aft propulsion system may provide vertical thrust during takeoff and landing. During flight phases in which the aircraft is in forward flight mode, the forward propulsion system may provide horizontal thrust, while the aft propulsion system propellers may be retracted in a fixed position to minimize drag. The aft electric propulsion system may be actively retracted while providing position monitoring. Some embodiments may include transitioning from vertical flight to horizontal flight and vice versa. In some embodiments, the transition may 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 may include a variable pitch mechanism that can change the collective angle of the forward propulsion system's propeller hub assembly blades for operation during hover, cruise, and transition phases. Some embodiments may include a conventional takeoff and landing (CTOL) configuration, with the tilter providing horizontal thrust for fixed-wing takeoff, cruise, and landing. The aft electric engine is not used to generate thrust during CTOL missions, and the aft propeller is stowed in place.

[0026] In some embodiments, the electric engines described herein may have design features to mitigate and prevent uncontained fires, such as having no nominal ignition source within the electric engine, utilizing a non-hazardous amount of flammable fluid contained in both the tilt engine and the lift engine with engine heating operating limits that may be more than 50°C below the auto-ignition temperature of the flammable fluid, overheat detection and protection, overvoltage detection and protection, and overcurrent detection and protection. In some embodiments, the design features of the electric engine may place the electric engine outside of a designated fire zone. In some embodiments, the flammable fluid may include oil, and the non-hazardous amount may be less than 1 quart, or 2 quarts, or 3 quarts, or 4 quarts, or 5 quarts, or 10 quarts, determined based on factors such as the size of the aircraft, the number of propellers, or the payload.

[0027] As disclosed herein, an electric engine may include an inverter and a motor, or an inverter, a gearbox, and a motor, in various configurations, such as the exemplary configurations described herein. For example, an electric engine may include an electric motor, a gearbox, and an inverter that all share the same central axis. Additionally, the central axis may be configured along the axis of the output shaft toward the aircraft's propeller. In such an exemplary configuration, the motor, gearbox, and inverter would all share the output shaft as the central axis and be oriented circularly around the output shaft. Additional embodiments may include a motor, gearbox, and inverter that are mounted together in an array or in a configuration in which some components, such as the motor and gearbox, are mounted together and other components, such as the inverter, are located elsewhere, but wiring is used to connect the electric engine.

[0028] As mentioned above, the electric engines for aircraft 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 motor's output shaft directly provides the speed and torque of the propeller shaft. Additional embodiments of the electric engine may include a motor, an inverter, and a gearbox, where the motor's output may be transmitted through a gearbox connected to the output shaft for the propeller, or a motor, an inverter, and a gearbox, where the output from the motor is decoupled from the propeller and transmitted through the gearbox, and the output shaft for the propeller re-enters the propeller through the gearbox and the motor. As described herein, the electric engine may contemplate any combination or orientation of some or all of the motor, inverter, and gearbox. Additionally, each configuration or orientation of the electric engine disclosed herein may include cooling via air cooling, a coolant, or a mixture of both.

[0029] For example, an electric engine configuration may include a motor and inverter, with the motor located between the aircraft propeller and the inverter. Additionally, the motor may include a gearbox. Furthermore, the inverter may share the same central axis as the motor, and the inverter may be located in an enclosure that cantilevers away from the rear of the motor and may be air-cooled. It is recognized that such an inverter orientation may not be an optimal configuration in terms of the enclosure required to achieve such a cantilevered orientation. Additionally, a motor in this configuration utilizing air cooling may include potting material, and air fins to assist in cooling the motor may result in a further significant increase in system mass.

[0030] Some embodiments may include an electric engine where the inverter module may be mounted outside the motor enclosure. Additional embodiments may include an electric engine where the inverter may be mounted above the electric motor such that the inverter's air cooling fins are below the propeller. Further embodiments may include the inverter mounted at the back of the motor with the air cooling fins facing radially outward, the inverter mounted at the front of the motor with the air cooling fins facing radially outward, the inverter mounted on the motor where the inverter is cooled by a liquid such as oil, or any other position of the inverter relative to the motor.

[0031] Embodiments of the electric motor may include a stator enclosure, a wound stator assembly, a rotor, various bearings, and any additional components that assist in transferring the speed and torque generated by the motor to the propeller.

[0032] It is understood that electric engines may generate heat during operation and may include thermal management systems to ensure that components of the electric engine do not fail during operation. In some embodiments, a coolant may be used and circulated throughout individual components of the engine, such as the inverter, gearbox, or motor, through some of the components, or through all of the engine's components, to help manage the heat present within the engine. Additional embodiments may include using air-cooling methods to cool the electric engine, or using a mixture of coolant and air to manage the heat generated by the electric engine during operation. In some embodiments, the coolant used may also be the same liquid used as a lubricant throughout the inverter, gearbox, or motor. For example, the inverter, gearbox, and motor may be cooled using liquid or air, or a mixture of air and liquid, such as using air cooling to cool the motor, using liquid cooling on the inverter and gearbox, or any other combination of air and liquid cooling throughout the inverter, gearbox, and motor, or even a subset of these components.

[0033] In some embodiments, oil may be used as a lubricant throughout the electric engine and as a coolant fluid to help manage heat generated by the engine during operation. Further to this example, various amounts of oil may be used to function as both a lubricant and a coolant fluid in the electric engine, such as less than one quart, less than two quarts, or any other amount of oil required to lubricate and cool the electric engine, with or without the assistance of air cooling. As disclosed herein, the electric engine may have different primary functions, such as being used only for ascent and landing and thus only in one orientation, or being used during all phases of flight, such as ascent, landing, and flight. An engine used during all phases of flight may experience various orientations throughout flight and may contain more lubricant and coolant than an engine used only in one orientation. As such, all engines on an aircraft may not need to contain the same amount of lubricant and coolant. For example, an ascent and landing engine may 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 exemplary embodiments referred to herein are representative and do not limit the amounts of lubricant and coolant that may be used in an electric engine.

[0034] It should be understood that using oil not only to lubricate but also to cool an electric engine rather than a separate coolant adds additional oil to the system, but eliminates a legacy component that could be used to cool such an electric engine. For example, if the electric engine were cooled by another liquid, such as glycol, the engine could include separate heat exchangers for both the lubricant fluid and the coolant fluid. Thus, in an embodiment where a single fluid is used for both lubrication and cooling, such as oil, there would be an increase in oil, but there would only be a need for one heat exchanger, so the overall system could have a reduced mass due to using fewer heat exchangers and potentially other components not being needed, and a more attractive drag profile. Furthermore, using one substance to lubricate and cool the engine could increase the efficiency of the system due to the reduced mass and the benefits of cooling the engine with a substance rather than relying on air cooling, which can be problematic to manage throughout the engine.

[0035] Additional embodiments of the electric engine may include various components to ensure that any flammable fluids are monitored and prevented from entering certain sections of the electric engine. Some embodiments may include an electric engine possessing a wet zone enclosure, which may be defined by the gearbox, motor, and / or heat exchanger. In some embodiments, the electric engine may have up to 4 liters or more of air in the motor gearbox enclosure in contact with the engine oil. For example, the electric engine may have up to 5 liters, 6 liters, 8 liters, 10 liters, or 20 liters of air in the motor gearbox enclosure in contact with the engine oil, based on factors such as aircraft size, number of propellers, or payload. Embodiments of the motor gearbox enclosure may use a breather to equalize internal and external pressure. Breather embodiments may include a breather that protrudes above a nearby design feature to prevent the intrusion of external fluids. Additional embodiments may include a breather with a screen and bypass entry path to prevent the intrusion of external debris. An embodiment may include sight glasses present on both the tilt and lift electric engines to ensure the oil is not overfilled or underfilled during maintenance.

[0036] Additional embodiments of the electric engine may include active protection features in the front and rear electric engines, such as monitoring internal temperatures throughout the engine, including oil temperature, stator winding sets, inverter bulk capacitors, power modules, control board power modules, control board control processors, control board monitor processors, internal hot spots, and various other locations throughout the engine. Embodiments may include overheating limits that take into account known fault temperatures and operating limits related to fluid autoignition temperatures. Some embodiments may include a high-voltage power system that may have a fuse at the high-voltage battery terminal that can quickly disconnect the engine electrical connection to irreversibly mitigate an overcurrent event. This overcurrent protection may be activated when the current draw of the electric engine is greater than the overcurrent operation. Thus, in some embodiments, a fault condition that leads to an overcurrent may only result in a temporary overheat, arc, or spark fault. 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 opened by an engine fuse. In some embodiments, the inverter detects AC overcurrents, isolates faulty phases, and / or continuously monitors the input DC voltage and applies protective actions to maintain the voltage below the overvoltage operating limit.

[0037] A. Exemplary Electric Aircraft Features FIG. 1 is an illustration of a perspective view of an exemplary VTOL aircraft consistent with disclosed embodiments. FIG. 2 is another illustration of a perspective view of an exemplary VTOL aircraft in an alternative configuration consistent with embodiments of the present disclosure. FIGS. 1 and 2 show VTOL aircraft 100, 200, respectively, in a cruise configuration and a vertical takeoff, landing, and hovering configuration (also referred to herein as a "climb" configuration), consistent with embodiments of the present disclosure. Corresponding elements in FIGS. 1 and 2 may bear like numerals and may refer to like elements of the aircraft 100, 200. The aircraft 100, 200 may include a fuselage 102, 202, wings 104, 204 attached to the fuselage 102, 202, and one or more aft stabilizers 106, 206 attached to the rear of the fuselage 102, 202. Multiple lift propellers 112, 212 may be mounted on the wings 104, 204 and configured to provide lift for vertical takeoff, landing, and hovering. Multiple tilt propellers 114, 214 may be mounted on the wings 104, 204 and tiltable between a climb configuration that provides a portion of the lift required for vertical takeoff, landing, and hovering, as shown in FIGURE 2, and a cruise configuration that provides forward thrust to the aircraft 100 for horizontal flight, as shown in FIGURE 1. As used herein, a tilt propeller climb configuration refers to any tilt propeller orientation in which the tilt propeller thrust is primarily providing lift to the aircraft, and a tilt propeller cruise configuration refers to any tilt propeller orientation in which the tilt propeller thrust is primarily providing forward thrust to the aircraft.

[0038] In some embodiments, the lift propellers 112, 212 may be configured to provide only lift, with all horizontal propulsion being provided by the tilt propellers. Thus, the lift propellers 112, 212 may be configured in a fixed position and may generate thrust only during the takeoff, landing, and hovering phases of flight. Meanwhile, the tilt propellers 114, 214 may be tilted upward into a lift configuration in which thrust from the propellers 114, 214 is directed downward to provide additional lift.

[0039] For forward flight, the tilt propellers 114, 214 may tilt from their climb configuration to their cruise configuration. In other words, the orientation of the tilt propellers 114, 214 may change from an orientation in which the tilt propeller thrust is directed downward (to provide lift during vertical takeoff, landing, and hovering) to an orientation in which the tilt propeller thrust is directed rearward (to provide forward thrust for the aircraft 100, 200). The tilt propeller assembly for a particular electric engine may tilt about an axis of rotation defined by the attachment point connecting the boom and the electric engine. When the aircraft 100, 200 is 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 that can be locked for cruising in a minimum-drag position, with one blade immediately ahead of the other, as illustrated in FIG. 1. In some embodiments, the lift propellers 112, 212 have three or more blades. In some embodiments, the tilt propellers 114, 214 may include more blades 116, 216 than the lift propellers 112, 212. For example, as illustrated in FIGS. 1 and 2, 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, the tilt propellers 114, 214 may each have two to five blades, or possibly more, depending on the design considerations and requirements of the aircraft.

[0040] 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 aft of the wings 104, 204, and at least a portion of the tilt propellers 114, 214 may be located forward of the wings 104, 204. In some embodiments, all of the lift propellers 112, 212 may be located aft of the wings 104, 204, and all of the tilt propellers 114, 214 may be located forward of the wings 104, 204. According to some embodiments, all of the lift propellers 112, 212 and tilt propellers 114, 214 may be mounted on the wings, i.e., the lift propellers or tilt propellers may not be mounted on the fuselage. In some embodiments, the lift propellers 112, 212 may all be located aft of the wings 104, 204, and the tilt propellers 114, 214 may all be located forward of the wings 104, 204. According to some embodiments, all of the lift propellers 112, 212 and tilt propellers 114, 214 may be positioned inboard of the ends of the wings 104, 204.

[0041] In some embodiments, the lift propellers 112, 212 and tilt propellers 114, 214 may be attached to the wings 104, 204 by booms 122, 222. The booms 122, 222 may be attached below the wings 104, 204, above the wings, and / or integrated into the wing profile. In some embodiments, the lift propellers 112, 212 and tilt propellers 114, 214 may be attached directly to the wings 104, 204. In some embodiments, one lift propeller 112, 212 and one tilt propeller 114, 214 may be attached to each boom 122, 222. The lift propellers 112, 212 may be attached to the aft end of the booms 122, 222, and the tilt propellers 114, 214 may be attached to the forward end of the booms 122, 222. In some embodiments, the lift propeller 112, 212 may be mounted in a fixed position on the boom 122, 222. In some embodiments, the tilt propeller 114, 214 may be mounted via a hinge to the forward end of the boom 122, 222. The tilt propeller 114, 214 may be mounted to the boom 122, 222 such that when in its cruise configuration, the tilt propeller 114, 214 is aligned with the body of the boom 122, 222 and forms a continuous extension of the forward end of the boom 122, 222 that minimizes drag for forward flight.

[0042] 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 on 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 wing 104, 204 may be 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.

[0043] In some embodiments, the aft stabilizer 106, 206 includes control surfaces such as one or more rudders, one or more elevators, and / or one or more combined rudder-elevators. The wing(s) may have any suitable design. In some embodiments, the wing has a tapered leading edge.

[0044] In some embodiments, the lift propeller 112, 212 or tilt propeller 114, 214 can tilt relative to at least one other lift propeller 112, 212 or tilt propeller 114, 214. As used herein, canting refers to the relative orientation of the lift / tilt propeller's axis of rotation about a line parallel to the longitudinal direction, similar to the roll degree of freedom of an aircraft. Tilting of the lift and / or tilt propeller can help minimize damage from propeller bursts and provide enhanced yaw control during flight by orienting the plane of rotation of the lift / tilt propeller disk (the blades and the hub to which they are attached) so as not to intersect critical parts of the aircraft (areas of the fuselage where personnel may be positioned, critical flight control systems, batteries, adjacent propellers, etc.) or other propeller disks.

[0045] FIG. 3 is an illustration of a top view of an exemplary VTOL aircraft consistent with embodiments of the present disclosure. The aircraft 300 shown in the figure may be a top view of the aircraft 100, 200 shown in FIGS. 1 and 2, respectively. As discussed herein, the aircraft 300 may include 12 electric propulsion systems distributed throughout the aircraft 300. In some embodiments, the distribution of the electric propulsion systems may include six forward electric propulsion systems 314 and six aft electric propulsion systems 312 mounted on forward and aft booms of the main wings 304 of the aircraft 300. In some embodiments, the length of the aft ends of the booms 324 from the wings 304 to the lift propellers may include similar aft ends of the length of the booms 324 across multiple aft ends of the booms. In some embodiments, the length of the aft ends of the booms may vary across the exemplary six aft ends of the booms. For example, each aft end of the booms 324 may have a different length from the wings 304 to the lift propeller, or a subset of the aft ends of the booms may be similar in length. In some embodiments, the forward end of the boom 322 may include various lengths from the wing 304 to the tilt propeller across the forward end of the boom. For example, as shown in FIG. 3 , the length of the forward end of the boom 322 from the tilt propeller closest to the fuselage to the wing 304 may include a longer length than the length of the forward end of the boom 322 from the wing 304 to the tilt propeller farthest from the fuselage. Some embodiments may include forward ends of the booms having similar lengths across the example six forward ends of the booms, or any other distribution of the lengths of the forward ends of the booms from the wing 304 to the tilt propeller. Some embodiments may include an aircraft 300 with eight electric propulsion systems, having four forward electric propulsion systems 314 and four aft electric propulsion systems 312, or any other distribution of forward and aft electric propulsion systems, including embodiments in which the number of forward electric propulsion systems 314 is less than or greater than the number of aft electric propulsion systems 312. Additionally, FIG. 3 illustrates an exemplary embodiment of a VTOL aircraft 300 with the forward propeller oriented horizontally for horizontal flight and the aft propeller blades 320 in a stowed position for forward flight.

[0046] As disclosed herein, the forward and aft electric propulsion systems may be clockwise (CW) or counterclockwise (CCW) types. Some embodiments may include a variety of forward electric propulsion systems having a mixture of both CW and CCW types. In some embodiments, the aft electric propulsion system may possess a mixture of CW and CCW types of systems among the aft electric propulsion systems.

[0047] FIG. 4 is a schematic diagram illustrating example propeller rotation for a VTOL aircraft consistent with disclosed embodiments. The aircraft 400 shown in the figure may be a top view of the aircraft 100, 200, and 300 shown in FIGS. 1, 2, and 3, respectively. The aircraft 400 may include six forward electric propulsion systems, with three of the forward electric propulsion systems being CW type 424 and the remaining three forward electric propulsion systems being CCW type. In some embodiments, the three aft electric propulsion systems may be CCW type 428 and the remaining three aft electric propulsion systems are CW type 430. Some embodiments may include the aircraft 400 with four forward electric propulsion systems and four aft electric propulsion systems, each having two CW type and two CCW type. In some embodiments, the propellers may counter-rotate relative to adjacent propellers to cancel torque steer generated by the propeller rotation and experienced by the aircraft fuselage or wings. In some embodiments, the difference in rotation direction may be achieved using engine rotation direction. In other embodiments, the engines may all rotate in the same direction and gearing may be used to achieve different propeller rotation directions.

[0048] Some embodiments may include an aircraft 400 possessing forward and aft electric propulsion systems, where the amount of CW type 424 and CCW type 426 is unequal between the forward electric propulsion systems, between the aft electric propulsion systems, or between the forward and aft electric propulsion systems.

[0049] FIG. 5 is a schematic diagram illustrating example power connections in a VTOL aircraft, consistent with disclosed embodiments. The VTOL aircraft may have various power systems connected to diagonally opposed electric propulsion systems. In some embodiments, the power systems may include high-voltage power systems. Some embodiments may include high-voltage power systems connected to electric engines via high-voltage channels. In some embodiments, the aircraft 500 may include six power systems 526, 528, 530, 532, 534, and 536 including batteries housed within wings 570 of the aircraft 500. In some embodiments, the aircraft 500 may include six forward electric propulsion systems having six electric engines 502, 504, 506, 508, 510, and 512 and six aft electric propulsion systems having six electric engines 514, 516, 518, 520, 522, and 524. In some embodiments, the batteries may be connected to the diagonally opposed electric engines. In such a configuration, first power system 526 may supply power to electric engine 502 via power connection channel 538 and may supply power to electric engine 524 via power connection channel 540. In some embodiments, first power system 526 may be paired with fourth power system 532 via power connection channel 542 that has a fuse to prevent excessive current from flowing through power systems 526 and 532. Further to this embodiment, VTOL air vehicle 500 may include second power system 528 that is paired with fifth power system 534 via power connection channel 548 that has a fuse, and may supply power to electric engines 510 and 516 via power connection channels 544 and 546, respectively. In some embodiments, third power system 530 may be paired with sixth power system 536 via fused power connection channel 554 and may supply power to electric engines 506 and 520 via power connection channels 550 and 552, respectively. Fourth power system 532 may also supply power to electric engines 508 and 518 via power connection channels 556 and 558, respectively. Fifth power system 534 may also supply power to electric engines 504 and 522 via power connection channels 560 and 562, respectively.Sixth power system 536 may also provide power to electric engines 512 and 514 via power connection channels 564 and 566, respectively.

[0050] As disclosed herein, an electric propulsion system may include an electric engine connected to a high-voltage power system, such as batteries located within the aircraft, via a high-voltage channel or power connection channel. Some embodiments may include various batteries housed within the aircraft wings with high-voltage channels going to the electric propulsion systems throughout the aircraft, including the wings and boom. In some embodiments, multiple high-voltage power systems may be used to create an electric propulsion system with multiple high-voltage power sources to avoid the risk of a single point of failure. In some embodiments, an aircraft may include multiple electric propulsion systems that may be wired to various batteries or power sources housed throughout the aircraft. It is recognized that such a configuration may be beneficial to avoid the risk of a single point of failure, where a failure of one battery or power source could result in a portion of the aircraft being unable to maintain the amount of thrust required to continue flight or perform a controlled landing. For example, if a VTOL possessed two forward electric propulsion systems and two aft electric propulsion systems, the forward electric propulsion system and the aft electric 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 forward and aft electric propulsion systems on opposite sides of the VTOL aircraft may remain operational, providing a more balanced flight or landing compared to failed forward and aft electric propulsion systems on the same side of the VTOL aircraft. Some embodiments may include four forward electric propulsion systems and four aft electric propulsion systems, with diagonally opposed electric engines connected to a common battery or power source. Some embodiments may include various configurations of electric engines electrically connected to the high-voltage power systems, such that in the event of a power failure, the risk of a single point of failure is avoided, and the flight phase in which the failure occurs may continue, or the aircraft may perform an alternate phase of flight in response to the failure.

[0051] 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 an electric propulsion system may include a high-voltage power grid that supplies high-voltage power to the electric engine and / or a low-voltage grid that supplies low-voltage DC power to the electric engine. Some embodiments may include the electric engine(s) in digital communication with a flight control system (“FCS”) that includes a flight control computer (“FCC”) that may send and receive signals to and from the electric engine including command and response data or status. Some embodiments may include the electric engine capable of receiving operating parameters from the FCC and communicating the operating parameters to the FCC, including speed, voltage, current, torque, temperature, vibration, propeller position, and any other value of the operating parameter.

[0052] In some embodiments, the flight control system may include a system that communicates with the electric engines and can send and receive analog / discrete signals to the electric engines to control devices that can redirect the thrust of the tilt propellers between a primarily vertical orientation in vertical flight mode and a primarily horizontal orientation in forward flight mode. In some embodiments, this system may be referred to as a tilt propeller system (“TPS”) and may be capable of communicating and directing additional features of the electric propulsion system.

[0053] FIG. 6 illustrates a block diagram of an example architecture and design of an electric propulsion unit 600 consistent with disclosed embodiments. In some embodiments, electric propulsion system 602 may include an electric engine subsystem 604 that may provide torque to a propeller subsystem 606 via a shaft to generate thrust for electric propulsion system 602. Some embodiments may include electric engine subsystem 604 receiving low-voltage DC (LV DC) power from a low-voltage system (LVS) 608. Some embodiments may include electric engine subsystem 604 receiving high-voltage (HV) power from a high-voltage power system (HVPS) 610 that includes at least one battery or other device capable of storing energy. In some embodiments, the high-voltage power system may include two or more batteries or other devices capable of storing energy and providing high-voltage power to electric engine subsystem 604. It is recognized that such a configuration may be advantageous in that failure of a single battery does not risk a single point of failure that could lead to failure of electric propulsion system 602.

[0054] Some embodiments may include an electric propulsion system 602 including an electric engine subsystem 604 that receives signals from and sends signals to a flight control system 612. In some embodiments, the flight control system 612 may include a flight control computer that can send commands to and receive status and data from the electric engine subsystem 604 using controller area network (“CAN”) data bus signals. While CAN data bus signals are used between the flight control computer and the electric engine(s), it should be understood that some embodiments may include any form of communication capable of sending and receiving data from the flight control computer to the electric engine(s). In some embodiments, the flight control system 612 may also include a tilt propeller system (“TPS”) 614 that can send and receive analog discrete data to and from the tilt propeller electric engine subsystem 604. The tilt propeller system 614 may include devices that communicate operating parameters to the electric engine subsystem 604 and that can articulate the orientation of the propeller subsystem 606 to redirect tilt propeller thrust during various phases of flight using mechanical means such as gearbox assemblies, linear actuators, and any other configuration of components for changing the orientation of the propeller subsystem 606.

[0055] As discussed throughout, exemplary VTOL aircraft may possess various types of electric propulsion systems, including tilt and lift propellers, including forward electric engines that have the ability to tilt during various phases of flight, and aft electric engines that may remain in one orientation and be active only during certain phases of flight (i.e., takeoff, landing, and hovering).

[0056] 7 is a schematic diagram illustrating an exemplary tilting electric propulsion system for a VTOL aircraft consistent with disclosed embodiments. The tiltable electric propulsion system 700 may include an electric engine assembly 702 aligned along a shaft 724 connected to an output shaft 738 mechanically coupled to a propeller assembly 720 comprising a hub, a spinner, and tilting propeller blades. In some embodiments, the electric engine assembly 702 may include a motor and gearbox assembly 704 aligned along and mechanically coupled to the shaft 724. In some embodiments, the motor and gearbox assembly 704 may include an electric motor assembly comprising a stator 706 and a rotor 708. As shown in FIG. 7 and in some embodiments, the stator 706 may include multiple stator windings connected to an inverter 716. In such a configuration, the stator 706 may incorporate one or more redundancies such that if one or more sets of windings fail, power can still be transferred to the stator 706 through one or more remaining windings, such that the electric engine assembly 702 retains power and continues to generate thrust at the propeller assembly 720.

[0057] In some embodiments, motor and gearbox assembly 704 may include a gearbox 710 aligned along shaft 724 to provide a gear reduction between torque on shaft 724 from an electric engine assembly including stator 706 and rotor 708 and output shaft 738. Torque applied to output shaft 738 may be transmitted to propeller assembly 720. Some embodiments may include gearbox 710 including an oil pump. In such embodiments, the oil pump may drive the circulation of oil throughout motor and gearbox assembly 704 at a speed equivalent to the rotation of output shaft 738 to cool and lubricate 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 output shaft 738. Some embodiments of motor and gearbox assembly 704 may include a propeller position sensor 712 present within the housing that may detect a magnetic field generated by the electric engine assembly to determine propeller position. Further embodiments may include a propeller position sensor 712 that is powered by the inverter 716 and transmits collected data to the inverter 716 .

[0058] In some embodiments, electric engine assembly 702 may also include an inverter assembly 714 substantially aligned along shaft 724. Inverter assembly 714 may include an inverter 716 and an inverter power supply 740. Inverter power supply 740 may receive low-voltage DC power from a low-voltage system 734 located outside electric engine assembly 702. Inverter power supply 740 may receive low-voltage DC power from a high-voltage power system 732 located outside electric engine assembly 702 that has been converted to low-voltage DC power via a DC-DC converter 742. Inverter 716 may supply high-voltage alternating current (AC) via at least one three-phase winding to a stator 706 of the electric engine assembly located within motor and gearbox assembly 704. Inverter assembly 714 may include inverter 716, which may receive flight control data from a flight control computing subsystem 736.

[0059] In some embodiments, motor and gearbox 704 may be located between inverter assembly 714 and propeller assembly 720. Some embodiments may also include a divider plate 744 coupled to motor and gearbox assembly 704 and inverter assembly 714. Divider plate 744 may create an enclosed environment for the upper part of motor and gearbox assembly 704 via the end bell assembly and for the lower part of inverter assembly 714 via the thermal plate. In some embodiments, divider plate 744 may serve as an integrated mounting bracket to support heat exchanger 718. Heat exchanger 718 may include, for example, folded fins or other types of heat exchangers. In some embodiments, electric propulsion system 700 may circulate oil or other coolant throughout electric engine assembly 702, motor and gearbox assembly 704, or inverter assembly 714 to transfer heat generated from the components to the oil or other coolant liquid. Heated oil or other coolant liquid may be circulated through the heat exchanger 718 to transfer heat to an airflow 722 passing through the fins of the heat exchanger.

[0060] In some embodiments, electric engine assembly 702 may be mounted or coupled to a boom structure 726 of the aircraft. Variable pitch mechanism 730 may be mechanically coupled to propeller assembly 720. In some embodiments, the variable pitch mechanism may abut electric engine assembly 702. In some embodiments, variable pitch mechanism 730 may be coupled to variable pitch mechanism 730 such that variable pitch mechanism 730 may be remotely mounted within a boom, wing, or fuselage of the aircraft. In some embodiments, variable pitch mechanism 730 may include a shaft or component that travels within or adjacent shaft 724 to propeller assembly 720. Variable pitch mechanism 730 may serve to change the collective angle of the propeller hub assembly blades of the forward electric engine as needed for operation during hover, transition, and cruise phases. Some embodiments may include electric engine assembly 702 mechanically coupled to a tilt propeller subsystem 728 that may redirect thrust between a primarily vertical direction during vertical flight mode and a primarily horizontal direction during forward flight mode. In some embodiments, the tilt propeller subsystem may abut a variable pitch mechanism 730. Some embodiments may include a tilt propeller subsystem 728 that includes various components located in various locations. For example, components of the tilt propeller subsystem may be coupled to the electric engine assembly 702, and other components may be coupled to the variable pitch mechanism 730. These various components of the tilt propeller subsystem 728 may work together to redirect the thrust of the tiltable electric propulsion system 700.

[0061] 8A-8C are illustrations of an exemplary tilt electric propulsion system for a VTOL aircraft consistent with disclosed embodiments. Figures 8A-8C have like numbers and refer to like elements of tiltable electric propulsion systems 800A, 800B, and 800C. As such, similar design considerations and configurations may be considered throughout the embodiments.

[0062] 8A and 8B illustrate side profile and perspective views, respectively, of tiltable electric propulsion systems 800A, 800B in a cruise configuration integrated with booms 812A, 812B consistent with the present disclosure. The tiltable propeller electric propulsion systems 800A, 800B may include electric engine assemblies 802A, 802B housed within the booms 812A, 812B of the VTOL aircraft. In some embodiments, the cruise configuration may include electric engine assemblies 802A, 802B located within the booms 812A, 812B. The electric engine assemblies 802A, 802B may include electric motor assemblies, gearbox assemblies, inverter assemblies with power connection channels 810A, 810B, and heat exchangers 804A, 804B, as described herein. The electric engine assemblies 802A, 802B may be mechanically coupled to propulsion assemblies 808A, 808B that include shaft flange assemblies 806A, 806B, spinners, and propeller blades.

[0063] 8C illustrates a top-down view along spinner 808C of tiltable electric propulsion system 800C in a raised configuration integrated with boom 812B consistent with the present disclosure. As shown in FIG. 8C, tiltable electric propulsion system 800C in the raised configuration may include electric engine assemblies 802A, 802B positioned outside of boom 812C and changing its orientation relative to boom 812C.

[0064] As discussed herein, an ascent electric propulsion system may be configured to provide thrust in one direction and may not provide thrust during all phases of flight. For example, an ascent system may provide thrust during takeoff, landing, and hovering, but not during cruise.

[0065] FIG. 9 is a schematic diagram illustrating an exemplary lift electric propulsion system for a VTOL aircraft consistent with disclosed embodiments. Lift electric propulsion system 900 may be mounted or coupled to a boom structure 924 of the aircraft. Lift electric propulsion system 900 may include an electric engine assembly 902 aligned along a shaft 940 connected to an output shaft 932 mechanically coupled to a propeller assembly 920 comprising hub and tilting propeller blades. In some embodiments, electric engine assembly 902 may include a motor and gearbox assembly housing 904 aligned along and mechanically coupled to shaft 940. In some embodiments, motor and gearbox assembly housing 904 may include an electric motor assembly comprising a stator 906 and a rotor 908. Stator 906 may include multiple stator windings connected to an inverter 916. In such a configuration, stator 906 may incorporate one or more redundancy and backup measures to avoid a single point of failure in case of an emergency. For example, stator 906 may include multiple windings such that if a winding fails, power can continue to be transferred to stator 906 through the remaining windings, allowing electric engine assembly 902 to maintain power and continue to generate thrust at propeller assembly 920.

[0066] In some embodiments, motor and gearbox assembly housing 904 may include a gearbox 910 aligned along shaft 940 to provide a gear reduction between torque on shaft 932 from an electric engine assembly comprising stator 906 and rotor 908 and output shaft 932. Torque applied to output shaft 932 may be transmitted to propeller assembly 920. Some embodiments may include gearbox 910 including a fluid pump for circulating a cooling and / or lubricating fluid. In the embodiment shown, the fluid pump is an oil pump. In such an embodiment, the oil pump may drive the circulation of oil throughout motor and gearbox assembly housing 904 at a speed equivalent to the rotation of output shaft 932 to cool and lubricate gearbox and electric motor components. Some embodiments of motor and gearbox assembly housing 904 may include a propeller position sensor 912 present within the housing that may detect a magnetic field generated by the electric engine assembly to determine propeller position. Further embodiments may include a propeller position sensor 912 that is powered by inverter 916 and transmits collected data to inverter 916, which may be forwarded to flight control computing system 930, among other flight control data.

[0067] In some embodiments, electric engine assembly 902 may also include an inverter assembly housing 914 aligned along an axis coextensive with the shaft axis. Inverter assembly housing 914 may include an inverter 916 and an inverter power supply 934. Inverter power supply 934 may receive low-voltage DC power from a low-voltage system 928 located outside electric engine assembly 902. Inverter power supply 934 may receive low-voltage DC power from a high-voltage power system 926 located outside electric engine assembly 902 that has been converted to low-voltage DC power via a DC-DC converter 936. Inverter 916 may provide high-voltage AC power via at least one three-phase winding to a stator 906 of the electric engine assembly located within motor and gearbox assembly housing 904. Inverter assembly 914 may include inverter 916, which may send data to and receive data from a flight control computing subsystem 930.

[0068] In some embodiments, motor and gearbox housing 904 may be located between inverter assembly housing 914 and propeller assembly 920. Some embodiments may include a divider plate 938 coupled to motor and gearbox assembly housing 904 and inverter assembly housing 914. Divider plate 938 may create an enclosed environment for the upper part of motor and gearbox assembly housing 904 via the end bell assembly and for the lower part of inverter assembly housing 914 via the thermal plate. In some embodiments, divider plate 938 may serve as an integrated mounting bracket to support heat exchanger 918. Heat exchanger 918 may include, for example, folded fins or other types of heat exchangers. In some embodiments, electric propulsion system 900 may circulate oil or other coolant liquid throughout electric engine assembly 902, motor and gearbox assembly 904, or inverter assembly 914 to transfer heat generated from the components to the oil or other coolant liquid. Heated oil or other coolant liquid may be circulated through the heat exchanger 918 to transfer heat to an airflow 922 passing over the fins of the heat exchanger.

[0069] In some embodiments, the tiltable electric propulsion system and the lift electric propulsion system may possess similar components. This may be advantageous with respect to many design considerations present in VTOL aircraft. For example, from a manufacturability perspective, different types of electric propulsion systems with similar components may be beneficial from a manufacturing efficiency perspective. Furthermore, having similar components may be beneficial from a risk management perspective because similar components have similar failure points that can be better identified and designed around when comparing a system with similar components to a system with different components and configurations.

[0070] It should be understood that while the tiltable electric propulsion system may have additional, and in some embodiments, different, components compared to the lift electric propulsion system, in some embodiments the tiltable electric propulsion system and the lift electric propulsion system may have the same configuration of components. For example, in some embodiments the tiltable and lift electric propulsion systems may include the same components, and the lift electric propulsion system may be coupled to the boom, wing, or fuselage of the aircraft such that it may not be able to provide thrust in as many directions as the tiltable electric propulsion system.

[0071] 10A-10B are illustrations of exemplary electric ascent propulsion systems for VTOL aircraft consistent with disclosed embodiments. 10A and 10B share similar numerals and refer to similar elements of electric ascent propulsion systems 1000A and 1000B. As such, similar design considerations and configurations may be considered throughout the embodiments.

[0072] 10A illustrates a side profile of an electric lift propulsion system 1000A in a lift configuration integrated with a boom 1010A consistent with the present disclosure. The electric lift propulsion system 1000A may comprise an electric engine assembly 1002A housed within the boom 1010A of the VTOL aircraft. In some embodiments, the lift configuration may include an electric engine assembly 1002A positioned vertically within the boom 1010A. The electric engine assembly 1002A may comprise an electric motor assembly, a gearbox assembly, an inverter assembly with a power connection channel 1008A, and a heat exchanger 1004A, as described herein. The electric engine assembly 1002A may be mechanically coupled to a propulsion assembly 1006A comprising a shaft flange assembly and propeller blades.

[0073] FIG. 10B illustrates a top-down view of an electric lift propulsion system 1000B in a lift configuration integrated with a boom 1010B, consistent with the present disclosure.

[0074] Some embodiments of the disclosed electric engine may 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 may be used and circulated throughout individual components of the engine, such as the inverter, gearbox, or motor, through some of the components, or through all of the engine's components, to help manage the heat present within 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 by the electric engine during operation. In some embodiments, the coolant used may also be the same liquid used as a lubricant throughout the inverter, gearbox, or motor. For example, the components of the electric engine may be cooled using liquid or air, or using a mixture of air and liquid cooling. As another example, the motor may be cooled using air cooling, and the inverter and gearbox may be cooled using liquid cooling. It should be understood that a mixture of cooling may be used for any combination of electric engine components or within each component.

[0075] In some embodiments, oil may be used as a lubricant throughout the electric engine and as a coolant fluid to help manage heat generated by the engine during operation. Further to this example, different amounts of oil may be used to function as both a lubricant and a coolant fluid within the electric engine, such as in combination with 1 quart or less, 1.5 quarts, 2 quarts, 2.5 quarts, 3 quarts, 5 quarts, or any other amount of oil needed to lubricate and cool the electric engine, with or without the assistance of air cooling. In some embodiments, the amount of oil or liquid used in the system in connection with cooling may be determined based on the amount of heat needed to drive heat transfer from the components of the electric propulsion system. As disclosed herein, electric engines may have different primary functions, such as being used only for ascent and landing and thus only in one orientation, or being used during all phases of flight, such as ascent, landing, and flight. An engine used during all phases of flight may experience various orientations throughout flight and may contain more lubricant and coolant than an engine used only in one orientation. As such, all engines on an aircraft may not contain the same amount of lubricant and coolant. For example, ascent and landing engines may require less than one quart of oil, while engines operating during all phases of flight may require more than one quart of oil. In some embodiments, the amount of cooling oil or liquid may be adequate to provide sufficient thermal mass to drive heat transfer from the components of the electric propulsion system, regardless of the orientation of the electric propulsion system. The embodiments discussed herein are exemplary and non-limiting and do not determine the limits on the amounts of lubricants and coolants that may be used in an electric engine.

[0076] Some embodiments may use oil to lubricate and cool the electric engine. Such embodiments may require an additional volume of oil. In such embodiments, the additional oil may allow for the elimination of conventional components that may be used to cool such an electric engine. For example, if the electric engine were cooled by another liquid, such as glycol, the engine may include separate heat exchangers for both the lubricant fluid and the coolant fluid. Thus, in embodiments in which a single fluid is used for both lubrication and cooling, such as oil, there would be an increase in oil, but only the need for one heat exchanger, so the overall system may have a reduced mass due to using fewer heat exchangers and potentially other components not being needed, and there may be a more attractive drag profile. Furthermore, using one substance to lubricate and cool the engine may increase the efficiency of the system due to the reduced mass and the benefits of cooling the engine with a substance rather than relying on air cooling, which can be problematic to manage throughout the engine.

[0077] Some embodiments of the electric engine may include various components for monitoring flammable fluids and preventing flammable materials from entering certain sections of the electric engine. Some embodiments may include an electric engine with a wet zone enclosure, which may be defined by the gearbox, motor, and / or heat exchanger. In some embodiments, the electric engine may have up to four liters or more of air within the motor gearbox housing, which is in contact with the engine oil. Some embodiments of the motor gearbox housing may use a breather to equalize internal and external pressure. Some embodiments of the breather may include a protrusion over a nearby design feature to prevent the inadvertent intrusion of external fluids. Some embodiments may include a breather with a screen and bypass entry path to prevent the intrusion of external debris. Some embodiments may include sight glasses present on both the tilt electric engine and the lift electric engine to ensure the oil is not overfilled or underfilled during maintenance.

[0078] Some embodiments of the electric engine may include active protection features in the forward and aft electric engines, as needed, such as monitoring vibrations throughout the engine and internal temperatures throughout the engine, such as oil temperature, stator winding setpoint temperature, inverter bulk capacitor temperature, power module temperature, control board power module temperature, control board control processor temperature, control board monitor processor temperature, internal hot spot temperature, and various other operating conditions. Such monitoring may be achieved using various sensors located throughout the electric propulsion system and the aircraft. Embodiments may include vibration limits based on known fault points or component resonances, and overheat limits set based on known fault temperatures and operating limits related to fluid autoignition temperatures. In some embodiments, the various sensors used to monitor operating conditions throughout the engine may report operating conditions to a flight control system. Some embodiments may include threshold operating values ​​that may be required before an operating value is transmitted to or flagged by the flight control system. In some embodiments, the flight control system may act to reduce the amount of power directed to the electric propulsion system in response to detecting an operating condition. Some embodiments may include reducing the amount of power to an electric propulsion system to reduce mechanical wear or friction sparks from vibrations and / or reducing power in an attempt to reduce the temperature of components present within the electric propulsion system. Additionally, some embodiments may include reducing power to an electric propulsion system whose detected inverter efficiency is below a target efficiency. In some embodiments, for example, if 12 electric propulsion systems are present in an aircraft, the flight control system may act to reduce or terminate power to a single electric propulsion system while increasing power directed to the remaining electric propulsion systems or a subset thereof to counteract a reduction in lift generated by one electric propulsion system. In some embodiments, the flight control system may establish various thresholds for operating conditions to correspond to a reduction or increase in power to the electric propulsion systems.

[0079] Some embodiments may include a high-voltage power system that may have a fuse at the high-voltage battery terminal that can quickly and irreversibly disconnect the engine electrical connection to mitigate and prevent an overcurrent event. Such overcurrent protection may be activated when the current draw of the electric engine is greater than the overcurrent operation. Thus, in some embodiments, a fault condition that leads to an overcurrent may result only in temporary overheating, arcing, or sparking. 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 opened by the engine fuse. In some embodiments, the inverter may detect AC overcurrents, isolate faulty phases, and / or continuously monitor the input DC voltage and apply protective action to maintain the voltage below the overvoltage operation limit.

[0080] During takeoff, landing, hovering, and cruising, the motors and associated control components of a VTOL aircraft may generate heat. The heat must be dissipated to prevent degradation or damage to the motors, control components, and other elements of the VTOL aircraft. In some types of VTOL aircraft, such as electric VTOL (eVTOL) aircraft, thermal control is also important, for example, to maintain optimal energy efficiency of battery-powered components.

[0081] Some elements may generate high heat loads only during certain periods of operation. For example, some lift propellers may be used only during takeoff, landing, and hovering and may be shut off during cruise. Thus, such lift propellers may generate high heat loads during takeoff, landing, and hovering, but generate little or no heat during cruise.

[0082] B. Exemplary Inverter Embodiments FIG. 11 illustrates a portion of an electric propulsion system 1100 for a vertical take-off and landing (VTOL) aircraft, consistent with some embodiments of the present disclosure. The electric propulsion system 1100 may provide a dual three-phase system for motor control. As shown in FIG. 11 , 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 to an alternating current (AC) voltage to drive a first set of stator windings of the electric motor M1 in response to a first pulse-width modulation (PWM) vector. The second inverter circuit 1120 is configured to convert the DC bus voltage Vbus on the bus of the second inverter circuit 1120 to an AC voltage to drive a second set of stator windings of the electric motor M1 in response to the second PWM vector. 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 may be less than or equal to 0.25%, 0.5%, 1%, or 2% of the switching cycle period. For example, the delay may be less than or equal to 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), where 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).

[0083] In particular, in electric propulsion system 1100, two inverter circuits 1110 and 1120 are electrically coupled to an internal high-voltage DC supply 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 dual three-phase motor M1. The dual inverter drive system shown in Figure 11 can improve motor performance and system reliability by increasing the number of phases.

[0084] As shown in the figure, inverter circuits 1110 and 1120 are each configured to convert a bus voltage Vbus on a high-voltage DC supply bus into three-phase AC power to drive motor M1. When inverter circuits 1110 and 1120 convert DC power to AC power, there is a voltage difference between the power source and the neutral point of the load, which is referred to as a common-mode voltage. Common-mode currents resulting from the common-mode voltage in the inverters can be harmful to electrical systems. Specifically, common-mode voltages can result in motor failure, premature bearing failure, controller glitches, and the like. In response to reduce common-mode noise, filter components can be installed in electric propulsion system 1100.

[0085] For example, electric propulsion system 1100 may include a DC common mode filter 1130 and AC common mode chokes 1140 and 1150. DC common mode filter 1130 may be coupled to bus capacitor 1170 and configured to reduce common mode signals on the DC sides of first inverter circuit 1110 and second inverter circuit 1120. One or more AC common mode chokes 1140 and 1150 may be coupled to the AC sides of first inverter circuit 1110 or second inverter circuit 1120 to reduce common mode signals.

[0086] For example, the DC common mode filter 1130 may be located between the DC power supply 1160 and the bus capacitor 1170 and may be formed by a set of DC-side chokes 1132 and a set of DC common mode filter capacitors 1134, 1136. The DC-side choke 1132 may be configured so that the positive and negative wires are wound around the same magnetic core. Thus, the DC-side choke 1132 and the DC common mode filter capacitors 1134, 1136 may be configured to reduce common mode signals on the DC side. However, a larger filter may add volume and mass, resulting in further losses.

[0087] In some embodiments, the electric propulsion system 1100 may achieve common-mode voltage cancellation by applying appropriate space vector modulation (SVM) to the inverter circuits 1110 and 1120.

[0088] The winding arrangement of the motor M1 employed in the electric propulsion system 1100 may vary in various embodiments. For example, the phase difference in electrical angle between the two sets of three-phase windings may 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 shifted by substantially 180 degrees (e.g., plus or minus 5 degrees). That is, the motor phasing between the independent winding sets may be out of phase by approximately 180 degrees.

[0089] In some embodiments, the first inverter circuit 1110 is controlled using standard midpoint-referenced space vector modulation (SVM), while the second inverter circuit 1120 is controlled using inverted midpoint-referenced space vector modulation. In such an operating mode, common-mode voltages can be counteracted via equal and opposite PWM vectors.

[0090] Figure 12A is a diagram illustrating PWM vectors for controlling the first inverter circuit 1110 and the second inverter circuit 1120 in the electric propulsion system 1100 of Figure 11, consistent with some embodiments of the present disclosure. Figure 12B is a diagram illustrating three-phase voltages u1, v1, and w1 output by the first inverter circuit 1110 and three-phase voltages u2, v2, and w2 output by the second inverter circuit 1120 over one period, consistent with some embodiments of the present disclosure.

[0091] In some embodiments, a space vector modulation (SVM) algorithm is applied to control pulse width modulation (PWM) to generate AC voltages from DC voltages and used to drive three-phase motors at varying speeds. As will be appreciated, different SVM algorithms may have different qualities and computational requirements. As shown in FIG. 12A, for a three-leg inverter using space vector modulation, there are eight possible switching vectors SV0-SV7. An exemplary PWM vector V for the first inverter circuit 1110 is INV 1 and an exemplary PWM vector V of the second inverter circuit 1120. INV 2 and are shown in FIG. 12A.

[0092] During operation, the switches in the inverter circuits 1110 and 1120 are controlled so that both switches in the same leg (i.e., the upper switch and the lower switch) are not turned on at the same time to avoid shorting the DC power supply. This can be achieved by complementary operation of the switches in the same leg. That is, for each output leg, when the upper switch is on, the lower switch is off, and vice versa. Therefore, the switching vectors SV0-SV7 include six active switching vectors SV1-SV6 and two zero vectors SV0 and SV7.

[0093] As shown in FIG. 12A , switching vector SV0={000} represents that the upper switches of the three phases U, V, and W are off, while the lower switches of the three phases U, V, and W are on. Switching vector SV1={100} represents that the upper switch of the U phase is on and the upper switches of the V and W phases are off. Switching vector SV2={110} represents that the upper switches of the U and V phases are on and the upper switch of the W phase is off. Switching vector SV3={010} represents that the upper switch of the V phase is on and the upper switches of the U and W phases are off. Switching vector SV4={011} represents that the upper switches of the V and W phases are on and the upper switch of the U phase is off. Switching vector SV5={001} represents that the upper switch of the W phase is on and the upper switches of the U and V phases are off. The switching vector SV6={101} indicates that the upper switches of the U and W phases are on and the upper switch of the V phase is off. The switching vector SV7={111} indicates that the upper switches of the three phases U, V, and W are on, while the lower switches of the three phases U, V, and W are off.

[0094] In some embodiments, the first inverter circuit 1110 is controlled according to a standard midpoint-based SVM starting from switching vector SV0 where the upper switches of the three phases U, V, and W are off, and the second inverter circuit 1120 is controlled according to an inverted midpoint-based SVM starting from switching vector SV7 where the upper switches of the three phases U, V, and W are on. Thus, as shown in FIG. 12B , at each stage of the operating period, the three-phase voltages u1, v1, and w1 output by the first inverter circuit 1110 are complementary to the three-phase voltages u2, v2, and w2 output by the second inverter circuit 1120, respectively. Thus, the resulting PWM vector V for the first inverter circuit 1110 is INV 1 and the PWM vector V for the second inverter circuit 1120. INV2 and 1110 are equal and opposite PWM vectors. Therefore, the resulting common-mode voltage V INV1-CM and the common mode voltage V generated from the second inverter circuit 1120. INV2-CM and have the same value but opposite signs, and due to the cancellation of the common-mode voltage, the total common-mode voltage V SYS-CM is zero.

[0095] In some embodiments, therefore, a six-phase machine may be implemented to achieve functionality similar to a three-phase machine while reducing the common-mode noise encountered by the system. In some embodiments, common-mode voltage cancellation between any number of phases may be implemented so that a combination of PWM vectors can cancel or reduce the common-mode noise encountered by the system. In some embodiments, the first and second inverter circuits 1110 and 1120 may be configured to draw power simultaneously from the power source (e.g., from the DC bus capacitor 1170), and the first and second inverter circuits 1110 and 1120 operate according to PWM signals with equal duty cycles but that are substantially 180 degrees out of phase.

[0096] The common-mode voltage cancellation achieved by the electric propulsion system 1100 with two inverters can reduce the volume and mass of the required electromagnetic compatibility (EMC) components in the system. In some embodiments, the common-mode voltage cancellation can provide at least 30-40 dB lower noise levels at low frequencies, which has a substantial impact on filter size. For example, the required size of the magnetic core for the DC-side choke can be significantly reduced, thereby reducing the overall weight of the engine.

[0097] 13A illustrates an inverter circuit 1300 for motor control in an electric propulsion system consistent with some embodiments of the present disclosure. Inverter circuit 1300, which is a voltage source inverter (VSI), may use FETs S1-S6 as switches. FETs S1-S6 may be selectively turned on or off in response to control signals from a controller according to the SVM algorithm described above to generate a three-phase AC output voltage for driving motor M1.

[0098] 13A, inverter circuit 1300 includes a bus capacitor 1320 configured to stabilize a direct current (DC) bus voltage Vbus and switches (e.g., FETs S1-S6) forming a plurality of phase legs, each of which includes an upper switch (e.g., FET S1, S3, or S5) disposed between the positive terminal of bus capacitor 1320 and the AC output terminal of the phase leg, and a lower switch (e.g., FET S2, S4, or S6) disposed between the negative terminal of bus capacitor 1320 and the AC output terminal of the phase leg.

[0099] In inverter circuit 1300, FETs S1-S6 may operate independently. During operation, if any of FETs S1-S6 malfunctions, a single-phase short-circuit fault occurs if one leg is shorted, which can result in uncontrolled current flow through the corresponding phase as the machine rotates. This type of fault could lead to a fire hazard, create drag on the system, add very high torque pulsations to the system, or any other potential hazard. Therefore, it would be advantageous to reduce the torque present in response to a single-phase short-circuit fault in system operation.

[0100] The inverter circuit 1300 includes a first discharge circuit 1310 between 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 with the discharge resistor 1312 to provide a discharge path for discharging energy stored in the bus capacitor 1320 (i.e., the bulk capacitor) when the switch 1314 is closed 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 at the appropriate time to remove the energy stored in the bus capacitor 1320.

[0101] In some embodiments, FETs S1-S6 may be further controlled to short out bus capacitor 1320 in response to the DC bus voltage Vbus being below a threshold fault condition. Alternatively, during discharge of bus capacitor 1320, when bus voltage Vbus is below a desired safety threshold, bridge short-through of all FETs S1-S6 may be applied to short out the HV bus to ensure a safe discharge process and distribute heat generated during a single-phase fault to all FETs S1-S6 in inverter circuit 1300 to ensure safety and reduce damage when a fault occurs.

[0102] 13A, pyrofuses F1 and F2 are coupled between the inverter circuit 1300 and a DC voltage source Vin (e.g., a battery pack). For example, pyrofuses F1 and F2 may be a type of fuse configured to be activated by an external source when circuit disconnection and isolation is required. For example, a first pyrofuse F1 may be coupled between the positive terminal of the DC voltage source Vin and the positive terminal of the bus capacitor 1320. A second pyrofuse F2 may be coupled between the negative terminal of the DC voltage source Vin and the negative terminal of the bus capacitor 1320.

[0103] 13B, which illustrates another inverter circuit 1300 for motor control in an electric propulsion system, consistent with certain embodiments of the present disclosure. The inverter circuit 1300 of FIG. 13B is also a voltage source inverter (VSI) that uses FETs S1-S6 as switches to output a three-phase AC output voltage to drive motor M1.

[0104] Compared to the inverter circuit 1300 of FIG. 13A, the inverter circuit 1300 of FIG. 13B further includes a comparator circuit 1330 and a second discharge circuit 1340 that form another discharge path across the bus capacitor 1320.

[0105] 13B , the comparator circuit 1330 is configured to monitor the bus voltage V across the bus capacitor 1320 when the inverter circuit 1300 discharges the bus capacitor 1320. In some other embodiments, the comparator circuit 1330 may also be configured to monitor the voltage across any other component(s) in the inverter circuit 1300 to determine whether to discharge energy through the second discharge circuit 1340 according to the voltage measurement. The second discharge circuit 1340 is configured to provide a rapid discharge path for any remaining energy stored in the bus capacitor 1320. In some embodiments, a threshold may be designed to confirm that a power source (e.g., a battery) has been disconnected from the HV DC bus.

[0106] For example, the comparator circuit 1330 may include a resistor divider 1332, a comparator 1334, and a logic circuit 1336. The resistor divider 1332 includes resistors R1 and R2 connected in series to provide a voltage V1. The comparator 1334 is configured to compare the voltage V1 output from the resistor divider 1332 with a reference voltage Vref. The second discharge circuit 1340 may include a rapid discharge component 1342 and a switch 1344 connected in series.

[0107] As the bus capacitor 1320 is discharged through the first discharge circuit 1310, the bus voltage V gradually decreases, and the voltage V1, which is a fixed fraction of the bus voltage V, also decreases. When the voltage V1 is lower than the reference voltage Vref, the bus voltage V is determined to be lower than a predetermined threshold, indicating that the power source and the HV DC bus are disconnected. The comparator 1334 is configured to output a corresponding signal (e.g., a logic 1 signal) in response to the voltage V1 being lower than the reference voltage Vref. The logic circuit 1336 may be an AND gate configured to receive the signal output by the comparator 1334 and a command signal Cmd from the control circuit. When the command signal Cmd and the signal output by the comparator 1334 are both logic 1, the logic circuit 1336 may output a corresponding control signal to turn on the switch 1344 in the second discharge circuit 1340. Therefore, the bus capacitor 1320 can be discharged through the second discharge circuit 1340 to achieve rapid discharge.

[0108] For example, the fast discharge component 1342 may be a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof, but the present disclosure is not limited thereto. The fast discharge component 1342 may also be any other component capable of achieving a fast discharge to remove remaining energy on the HV DC bus.

[0109] In an alternative approach, in accordance with the circuit operation described above, inverter circuit 1300 may achieve rapid high-voltage discharge by monitoring the voltage across components in inverter circuit 1300 (e.g., bus capacitors) and performing short-through across FETs S1-S6 in inverter circuit 1300. In some embodiments, after initiating a first discharge phase to begin voltage discharge, the voltage across the bus capacitor is monitored using a comparator circuit until a threshold is reached. In response to the monitored voltage detection dropping to the threshold, a second discharge phase may be engaged in parallel. The second discharge phase may utilize a discharge component (e.g., rapid discharge component 1342), such as a TVS or MOV, to discharge the remaining voltage accordingly. In some other embodiments, threshold voltage detection may not be the only condition for initiating the second discharge phase. For example, a corresponding command (e.g., command signal Cmd) to initiate the second discharge phase may be generated and sent to one or more components (e.g., logic circuit 1336) for initiating the second discharge stage upon detection of the monitored voltage dropping to a threshold. In various embodiments, the particular threshold voltage may depend on the battery characteristics, resistor characteristics, the length of the discharge time, or any other components involved in the circuitry.

[0110] 14 illustrates an example flowchart of a method 1400 for controlling the inverter circuit 1300 of FIG. 13A or 13B to mitigate a single-phase short-circuit fault, consistent with some embodiments of the present disclosure. As shown in FIG. 14 , the method 1400 may include steps 1410, 1420, 1430, 1440, and 1450.

[0111] In step 1410, it is detected whether a fault has occurred in one of the switches (e.g., FETs S1-S6) in inverter circuit 1300. In some embodiments, inverter circuit 1300 is configured to determine whether a single-phase short-circuit fault has occurred in any one of FETs S1-S6. For example, inverter circuit 1300 may include a voltage or current sensing circuit or component for detecting a single-phase short-circuit fault based on a voltage signal or a current signal within inverter circuit 1300.

[0112] In step 1420, in response to detecting a single-phase short-circuit fault, a three-phase short is applied by controlling the low-side switch (e.g., FETs S2, S4, and S6 of FIG. 13B) or the high-side switch (e.g., FETs S1, S3, and S5 of FIG. 13B). In some embodiments, in response to determining that a single-phase short-circuit fault has occurred, in step 1420, inverter circuit 1300 is configured to apply a three-phase short to the healthy side of inverter circuit 1300. For example, if a low-side FET (e.g., S2, S4, or S6) is damaged, causing a single-phase short, the high-side FETs (e.g., S1, S3, and S5) may be controlled to apply a three-phase short, or vice versa.

[0113] After the three-phase short circuit is applied, in step 1430, the inverter circuit 1300 is disconnected from the power source in response to detecting a single-phase short circuit fault. In some embodiments, the inverter circuit 1300 is configured to instruct a battery management system (BMS) in the system to enable a protection mechanism to open the HV circuit from the faulty inverter circuit 1300. In some embodiments, the battery management system is housed in an HV junction box (HVJB) and is configured to monitor voltage, temperature, current, and insulation resistance, and to control pack contactors and pyrofuses to protect against fault conditions for safe operation.

[0114] In some embodiments, step 1430 includes steps 1432, 1434, and 1436. For example, step 1432 may send instructions from inverter circuit 1300 to the BMS. Then, in step 1434, the BMS may send one or more command signals to one or more pyrofuse drivers to activate one or more pyrofuses to disconnect inverter circuit 1300 from the power source.

[0115] In an alternative approach, in response to an instruction or command from the inverter circuit 1300, the BMS is configured to generate a pyro event that activates one or both of the pyro fuses F1 and F2 to disconnect the HV DC voltage source Vin from the faulty inverter circuit 1300. Additionally, in step 1436, the BMS may further transmit a confirmation signal to the inverter circuit 1300 after activating one or more pyro fuses to confirm that the pyro has been activated.

[0116] That is, the BMS is used to activate pyrofuses F1 and / or F2 when a short-circuit event occurs. In some embodiments, the BMS may also send a command signal to a corresponding pyrofuse driver to activate the pyrofuse, thus electrically isolating the battery pack from the connected inverter circuit 1300 when other types of faults occur, such as those to protect against overcurrent. In response to receiving a confirmation signal from the BMS, the inverter circuit 1300 may then confirm that the HV DC voltage source Vin (e.g., the battery pack) has been disconnected from the inverter circuit 1300.

[0117] In step 1440, the bus voltage Vbus across the bus capacitor 1320 of the inverter circuit 1300 is discharged after the inverter circuit 1300 is disconnected from the power source. In some embodiments, after receiving the confirmation 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 to form a discharge path for discharging the energy using the discharge resistor 1312. Thus, the bus voltage Vbus across the bus capacitor 1320 gradually decreases in the discharge process. In some embodiments, in step 1440, the inverter circuit 1300 may perform rapid discharge by using multiple discharge circuits in parallel.

[0118] For example, in step 1442, in response to determining that the inverter circuit 1300 is disconnected from the power source, a first switch (e.g., switch 1314 in FIG. 13B) in a first discharge circuit (e.g., first discharge circuit 1310 in FIG. 13B) is closed to provide a first discharge path. In step 1444, the bus voltage V across the bus capacitor 1320 is monitored by a comparator circuit (e.g., comparator circuit 1330 in FIG. 13B). In step 1446, in response to the bus voltage V being lower than a second threshold, a second switch (e.g., switch 1344 in FIG. 13B) in a second discharge circuit (e.g., second discharge circuit 1340 in FIG. 13B) is closed to provide a second discharge path in parallel to the first discharge path.

[0119] In step 1450, in response to the bus voltage Vbus being lower than a first threshold, multiple switches (e.g., FETs S1-S6 in FIG. 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. When the bus voltage Vbus is lower than a certain threshold voltage (e.g., approximately 50 V), the inverter circuit 1300 may apply a bridge short-through to short-circuit all six FETs S1-S6 to short-circuit the high-voltage bus. The threshold voltage may be designed based on practical needs. For example, the threshold voltage may range from approximately 40 V to 60 V. Because the short-circuit configuration is enabled after the bus voltage Vbus drops below a certain level during the rapid discharge process, the resulting current will be within safe current limits 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.

[0120] Through the operation of method 1400 described above, heat generated during a single-phase fault may be distributed among all FETs in inverter circuit 1300, ensuring safety and reducing damage when a fault occurs. Furthermore, any one or more of steps 1410-1450 performed in method 1400 may occur with a specific time frame between steps, such as nanoseconds, milliseconds, or any other time value. In some embodiments, a control circuit in the system may provide control signals to selectively open or close switches (e.g., FETs) in inverter circuit 1300. The control circuit may also provide pyro signals to activate one or both of pyrofuses F1 and F2. In some other embodiments, the system may also include multiple control circuits for sending pyro signals to activate pyrofuses F1 and F2 and control signals to control one or more of the FETs in inverter circuit 1300. It should be understood that any step(s) described herein are not necessarily performed in a specific number of stages or in a specific stage. Step(s) may be performed throughout method 1400 as needed. Additionally, while the above illustrations include example steps and / or operations, operations may be added, substituted, reordered, and / or eliminated as appropriate without departing from the spirit and scope of the present disclosure. It will be appreciated that the methods for controlling inverter circuits and inverter circuits disclosed in various embodiments may also be utilized in various fields or systems, including, but not limited to, fields such as automobiles, hybrid and electric vehicles, electric motors, etc.

[0121] Various embodiments herein are described in one aspect in the general context of method steps or processes that may be implemented by an integrated circuit that includes circuitry for performing the method for controlling an inverter circuit. The circuitry may be configured to perform the method steps or processes. 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.

[0122] Various embodiments herein are described, in one aspect, in the general context of method steps or processes that may be implemented by a computer program product embodied in a transitory or non-transitory computer-readable medium storing computer-executable instructions, such as program code, for execution by one or more processors or one or more controllers in a system. The computer-readable medium may include removable and non-removable storage devices, including, but not limited to, read-only memory devices (ROM), random-access memory devices (RAM), compact discs (CDs), digital versatile discs (DVDs), and the like.

[0123] Generally, program modules may include routines, programs, objects, components, data structures, etc. that perform particular tasks or implement particular abstract data types. Computer-executable instructions, associated data structures, and program modules represent examples of program code for executing steps of the methods disclosed herein. The particular sequence of such executable instructions or associated data structures represents examples of corresponding acts for implementing the functions described in such steps or processes.

[0124] 15, which is a graph illustrating bus voltage versus time for an HV DC bus during a discharge period consistent with certain embodiments of the present disclosure. As shown in curve 1510, in a conventional discharge process, the voltage, current, and charge may decay exponentially during capacitor discharge. However, as shown in curve 1520, certain embodiments of the present disclosure may enable rapid discharge. Thus, initially, the bus voltage V may decay exponentially along an exponential decay curve, but only after the bus voltage V reaches a certain threshold V at time T. TAs described in the above embodiment, when the inverter circuit 1300 detects that the bus voltage Vbus falls below a certain threshold V at time T, the bus voltage Vbus may rapidly fall to zero rather than continuing along an exponential decay curve. T Upon detecting that the voltage is lower than 1342, various circuitry (eg, fast discharge component 1342 in FIG. 13B) may be used to quickly discharge the remaining voltage.

[0125] Reference is made to FIGS. 16A-16D, which are graphs illustrating EMI noise consistent with certain embodiments of the present disclosure. FIGS. 16A-16B are graphs illustrating EMI noise detected on the positive and negative sides, respectively, of a dual-inverter system consistent with certain embodiments of the present disclosure. FIGS. 16C-16D are graphs illustrating EMI noise detected on the positive and negative sides, respectively, of a single-inverter system consistent with certain embodiments of the present disclosure. As shown in FIGS. 16A-16D, in various embodiments disclosed herein, the EMI noise in the single-inverter system is approximately 30-40 dB higher than the EMI noise in the proposed system using two inverters with PWM signals interleaved by substantially 180 degrees. Therefore, the common-mode noise suppression / cancellation achieved by an electric propulsion system having two inverters can reduce the volume and mass of required electromagnetic compatibility (EMC) components. As a result, a lighter-weight engine design can be realized.

[0126] In the foregoing specification, embodiments have been described with reference to numerous specific details that may vary from implementation to implementation. Certain adaptations and modifications of the described embodiments may be made. Also, it is intended that the sequence of steps depicted in the figures is for illustrative purposes only and is not intended to be limited to any particular sequence of steps. Thus, one skilled in the art will appreciate that these steps may be performed in different orders while implementing the same method.

[0127] As used herein, unless otherwise stated, the term "or" includes all possible combinations unless infeasible. For example, if it is stated that a module may include A or B, then the module may include A, or B, or A and B, unless otherwise stated or infeasible. As a second example, if it is stated that a module may include A, B, or C, then the module may include A, or B, or C, or A and B, or A and C, or B and C, or A, B, and C, unless otherwise stated or infeasible.

[0128] In the drawings and specification, exemplary embodiments are disclosed. Various modifications and variations can be made to the disclosed devices, systems, and related methods, and will be apparent to those skilled in the art. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed devices, systems, and related methods. It is intended that the specification and examples be considered as exemplary only, with the true scope being indicated by the following claims and their equivalents.

[0129] The embodiments may be further described using the following clauses. Clause Set 1 1. A propulsion system for an aircraft, 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 a 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; a second inverter circuit coupled to the capacitor and configured to convert a 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, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors. 2. A propulsion system as described in clause 1, wherein the first set of stator windings and the second set of stator windings are shifted by substantially 180 degrees. 3. A propulsion system as described in clause 1 or 2, wherein the first inverter circuit is controlled using midpoint reference space vector modulation. 4. A propulsion system as described in any one of the preceding clauses, wherein the second inverter circuit is controlled using inverted midpoint reference space vector modulation. 5. A propulsion system as described in 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. 6. The propulsion system of clause 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. A propulsion system as described in any one of the preceding clauses, further comprising 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. 8. A propulsion system as described in any one of the preceding clauses, further comprising one or more AC common mode chokes coupled to the AC side of the first inverter circuit or the second inverter circuit to reduce common mode signals. Clause Set 2 9. A method for controlling a propulsion system for an aircraft, comprising: Stabilizing the direct current (DC) bus voltage with a capacitor; converting the DC bus voltage to an alternating current (AC) voltage according to a first pulse width modulation (PWM) vector with a first inverter circuit coupled to a capacitor to drive a first set of stator windings of the electric motor; converting the DC bus voltage to an AC voltage in response to a second PWM vector 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; driving one or more propellers of the aircraft with an electric motor. 10. The method of clause 9, wherein the first set of stator windings and the second set of stator windings are shifted by substantially 180 degrees. 11. The method of clause 9 or 10, further comprising controlling the first inverter circuit using midpoint reference space vector modulation. 12. The method of any one of clauses 9-11, further comprising controlling the second inverter circuit using inverted midpoint reference space vector modulation. 13. Outputting a first set of three-phase AC voltages by a first inverter circuit to drive a first set of stator windings; 13. The method of any one of clauses 9-12, further comprising: outputting a second set of three-phase AC voltages by a second inverter circuit to drive a second set of stator windings. 14. The method of clause 13, 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. 15. The method of any one of clauses 9 to 14, further comprising reducing common mode signals on the DC sides of the first inverter circuit and the second inverter circuit with a DC common mode filter coupled to a capacitor. 16. The method of any one of clauses 9 to 15, further comprising reducing 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 sides of the first inverter circuit or the second inverter circuit. Clause Set 3 17. An integrated circuit comprising circuitry for performing a method for controlling a propulsion system for an aircraft, the circuitry comprising: controlling a first inverter circuit coupled to the capacitor to convert the DC bus voltage into an alternating current (AC) voltage to drive a first set of stator windings of the electric motor according to a first pulse width modulation (PWM) vector; and an integrated circuit configured to control a second inverter circuit coupled to the capacitor in response to the second PWM vector to convert the DC bus voltage to an AC voltage to drive a second set of stator windings of the electric motor, the first PWM vector and the second PWM vector being substantially equal and opposite vectors for driving one or more propellers of the aircraft by the electric motor. 18. The integrated circuit of clause 17, wherein the first set of stator windings and the second set of stator windings are shifted by substantially 180 degrees. 19. The integrated circuit of clause 17 or 18, wherein the circuitry is further configured to control the first inverter circuit using midpoint reference space vector modulation. 20. The integrated circuit of any one of clauses 17-19, wherein the circuitry is further configured to control the second inverter circuit using inverted midpoint reference space vector modulation. 21. The circuitry further comprises: controlling the first inverter circuit to output a first set of three-phase AC voltages by the first inverter circuit to drive the first set of stator windings; 21. The integrated circuit of any one of clauses 17 to 20, configured to control a second inverter circuit to output a second set of three-phase AC voltages by the second inverter circuit to drive a second set of stator windings. 22. The integrated circuit of clause 21, 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. 23. An integrated circuit according to any one of clauses 17 to 22, wherein a DC common mode filter is coupled to the capacitor and configured to reduce common mode signals on the DC side of the first inverter circuit and the second inverter circuit. 24. An integrated circuit according to any one of clauses 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 common mode signals. Clause Set 4 25. An inverter circuit for a propulsion system for an aircraft, comprising: a capacitor configured to stabilize a direct current (DC) bus voltage; a plurality of switches forming a plurality of phase legs, at least one of the phase legs including an upper switch disposed between a positive terminal of a 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; An inverter circuit, wherein the plurality of switches are controlled to short out the capacitor in response to the DC bus voltage being below a first threshold in a fault condition associated with the inverter circuit. 26. The inverter circuit of clause 25, wherein the first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and wherein 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 of clause 25 or 26, further comprising 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. 28. The inverter circuit of clause 27, wherein the second discharge circuit comprises a discharge element and a second switch connected in series. 29. The inverter circuit of clause 28, wherein the discharge element comprises a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof. 30. The inverter circuit of any one of clauses 25 to 29, 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. 31. A comparator circuit is a resistor divider comprising resistors connected in series to provide a first voltage, the first voltage being a fixed fraction of the DC bus voltage; a comparator coupled to the resistor voltage divider; and comparing the first voltage output from the resistor divider with a reference voltage; 31. The inverter circuit of clause 30, comprising: a comparator configured to output an output signal in response to the first voltage being lower than the reference voltage. 32. A comparator circuit is a logic circuit coupled to the comparator; and receiving an output signal from the comparator and a command signal from the control circuit; 32. The inverter circuit of claim 31, further comprising a logic circuit configured to output a control signal for selectively turning on a second switch of the second discharge circuit according to the output signal and the command signal. Clause Set 5 33. A method for controlling an inverter circuit, comprising: Detecting whether a fault occurs in one of a plurality of switches in the inverter circuit; disconnecting the inverter circuit from the power source in response to detecting a single phase short circuit fault; a first discharge circuit providing a first discharge path for discharging the DC bus voltage across a capacitor of the inverter circuit after the inverter circuit is disconnected from the power source; and controlling a plurality of switches in an inverter circuit to short out a capacitor in response to the DC bus voltage being less than a first threshold. 34. The method of clause 33, wherein the first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and the method includes discharging the capacitor through the inverter circuit by closing the first switch in response to a single-phase short circuit. 35. The method of clause 33 or 34, further comprising providing a second discharge path for discharging the DC bus voltage by a second discharge circuit coupled in parallel to the capacitor. 36. The inverter circuit of clause 35, wherein the second discharge circuit comprises a discharge element and a second switch connected in series. 37. The inverter circuit of clause 36, wherein the discharge element comprises a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof. 38. The method of any one of clauses 33-37, further comprising monitoring the DC bus voltage across the capacitor with a comparator circuit to determine whether to discharge the energy stored in the capacitor through a second discharge circuit. 39. Providing a first voltage by a resistor divider comprising resistors connected in series, the first voltage being a fixed fraction of the bus voltage; 39. The method of any one of clauses 33 to 38, further comprising: comparing the first voltage with a reference voltage by a comparator; and outputting an output signal in response to the first voltage being lower than the reference voltage. 40. receiving, by a logic circuit, an output signal from the comparator and a command signal from the control circuit; 40. The method of claim 39, further comprising: outputting, by the logic circuit, a control signal for selectively turning on a second switch in the second discharge circuit according to the output signal and the command signal. Clause Set 6 41. An integrated circuit comprising circuitry for implementing a method for controlling an inverter circuit, the circuitry comprising: Detecting whether a fault occurs in one of a plurality of switches in the inverter circuit; disconnecting the inverter circuit from the power source in response to detecting a single phase short circuit fault; controlling the inverter circuit to provide, with a first discharge circuit, a first discharge path for discharging the DC bus voltage across a capacitor of the inverter circuit after the inverter circuit is disconnected from the power source; and controlling an inverter circuit to control a plurality of switches in the inverter circuit to short out a capacitor in response to the DC bus voltage being lower than a first threshold. 42. The integrated circuit of clause 41, wherein the first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and wherein in response to a single-phase short circuit, the circuitry is configured to control the inverter circuit to discharge the capacitor by closing the first switch. 43. The circuit mechanism is 43. The integrated circuit of claim 41 or 42, configured to implement controlling the inverter circuit to provide a second discharge path for discharging the DC bus voltage by a second discharge circuit coupled in parallel to the capacitor. 44. The integrated circuit of clause 43, wherein the second discharge circuit comprises a discharge element and a second switch connected in series. 45. The integrated circuit of clause 44, wherein the discharge element comprises a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof. 46. ​​An integrated circuit as described in any one of clauses 43 to 45, wherein a comparator circuit monitors the DC bus voltage across the capacitor to determine whether to discharge the energy stored in the capacitor through a second discharge circuit. 47. Suppose the DC bus voltage across a capacitor is... Ch. providing a first voltage by a resistor divider comprising resistors connected in series, the first voltage being a fixed fraction of the bus voltage; 47. The integrated circuit of any one of clauses 41 to 46, wherein the integrated circuit is monitored by: comparing a first voltage with a reference voltage using a comparator; and outputting an output signal in response to the first voltage being lower than the reference voltage. 48. Suppose the DC bus voltage across a capacitor is... Ch. receiving, by a logic circuit, an output signal from the comparator and a command signal from the control circuit; and outputting a control signal for selectively turning on a second switch in the second discharge circuit according to the output signal and the command signal by the logic circuit. Clause Set 7 49. An inverter circuit for a propulsion system for an aircraft, comprising: a capacitor configured to stabilize a direct current (DC) bus voltage; a plurality of switches forming a plurality of phase legs, at least one of the phase legs including an upper switch disposed between a positive terminal of a 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; 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 lower than a threshold in a fault condition associated with the inverter circuit. 50. The inverter circuit of clause 49, wherein the first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and wherein in response to a single-phase short circuit, the inverter circuit is configured to discharge the capacitor by closing the first switch. 51. An inverter circuit as described in clause 49 or 50, wherein the second discharge circuit comprises a discharge element and a second switch connected in series. 52. The inverter circuit of clause 51, wherein the discharge element comprises a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof. 53. The inverter circuit of any one of clauses 49 to 52, 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. 54. A comparator circuit is a resistor divider comprising resistors connected in series to provide a first voltage, the first voltage being a fixed fraction of the DC bus voltage; a comparator coupled to the resistor voltage divider; and comparing the first voltage output from the resistor divider with a reference voltage; 54. The inverter circuit of claim 53, comprising: a comparator configured to output an output signal in response to the first voltage being lower than the reference voltage. 55. A comparator circuit is a logic circuit coupled to the comparator; and receiving an output signal from the comparator and a command signal from the control circuit; 55. The inverter circuit of claim 54, further comprising a logic circuit configured to output a control signal for selectively turning on a second switch of the second discharge circuit according to the output signal and the command signal. Clause Set 8 56. A method for controlling an inverter circuit, comprising: Detecting whether a fault occurs in one of a plurality of switches in the inverter circuit; disconnecting the inverter circuit from the power source in response to detecting a single phase short circuit fault; discharging a bus voltage across a bus capacitor of the inverter circuit after the inverter circuit is disconnected from the power source; providing a first discharge path using a first discharge circuit in response to determining that the inverter circuit is disconnected from the power source; and providing a second discharge path in parallel to the first discharge path using a second discharge circuit in response to the bus voltage being below a threshold. 57. The first discharge circuit includes a discharge resistor and a first switch connected in series with the discharge resistor, and discharges the bus voltage across the capacitor. 57. The method of clause 56, further comprising: in response to determining that the inverter circuit is disconnected from the power source, closing a first switch in the first discharge circuit to provide a first discharge path. 58. The second discharge circuit includes a discharge element and a second switch connected in series to discharge the bus voltage across the capacitor; 58. The method of claim 56 or 57, further comprising: in response to the bus voltage being lower than the threshold, closing a second switch in the second discharge circuit to provide a second discharge path in parallel with the first discharge path. 59. The method of clause 58, wherein the discharge element comprises a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof. 60. Discharging the bus voltage across a capacitor is 60. The method of any one of clauses 56-59, further comprising monitoring the bus voltage across the capacitor with a comparator circuit to determine whether to discharge the energy stored in the capacitor through a second discharge circuit. 61. Discharging the bus voltage across a capacitor is providing a first voltage by a resistor divider comprising resistors connected in series, the first voltage being a fixed fraction of the bus voltage; 61. The method of any one of clauses 56 to 60, further comprising: comparing the first voltage with a reference voltage by a comparator; and outputting an output signal in response to the first voltage being lower than the reference voltage. 62. Discharging the bus voltage across a capacitor is receiving, by a logic circuit, an output signal from the comparator and a command signal from the control circuit; 62. The method of claim 61, further comprising: outputting, by the logic circuit, a control signal for selectively turning on a second switch in the second discharge circuit according to the output signal and the command signal. Clause Set 9 63. An integrated circuit comprising circuitry for implementing a method for controlling an inverter circuit, the circuitry comprising: Detecting whether a fault has occurred in one of the plurality of switches in the inverter circuit; disconnecting the inverter circuit from the power source in response to detecting a single-phase short-circuit fault; providing a first discharge path using a first discharge circuit in response to determining that the inverter circuit is disconnected from the power source; and in response to the bus voltage being lower than a threshold, providing a second discharge path in parallel to the first discharge path using a second discharge circuit, thereby controlling the inverter circuit to discharge the bus voltage across a capacitor of the inverter circuit after the inverter circuit is disconnected from the power source. 64. The first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and the inverter circuit comprises: 64. The integrated circuit of claim 63, wherein the integrated circuit is controlled to discharge the bus voltage across the capacitor by closing a first switch in a first discharge circuit to provide a first discharge path in response to determining that the inverter circuit is disconnected from the power source. 65. The second discharge circuit includes a discharge element and a second switch connected in series, and the inverter circuit includes: 65. The integrated circuit of clause 63 or 64, wherein the integrated circuit is controlled to discharge the bus voltage across the capacitor in response to the bus voltage being lower than a threshold by closing a second switch in a second discharge circuit to provide a second discharge path. 66. The method of clause 65, wherein the discharge element comprises a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof. 67. Suppose the bus voltage across a capacitor is 67. The method of any one of clauses 63 to 66, wherein the capacitor is discharged by monitoring the bus voltage across the capacitor with a comparator circuit to determine whether to discharge the energy stored in the capacitor through a second discharge circuit. 68. If the bus voltage across a capacitor is providing a first voltage by a resistor divider comprising resistors connected in series, the first voltage being a fixed fraction of the bus voltage; An integrated circuit as described in any one of clauses 63 to 67, which is discharged by comparing a first voltage with a reference voltage using a comparator, and outputting an output signal in response to the first voltage being lower than the reference voltage. 69. If the bus voltage across a capacitor is receiving, by a logic circuit, an output signal from the comparator and a command signal from the control circuit; The integrated circuit of any one of clauses 63 to 68, wherein the discharge is performed by outputting a control signal for selectively turning on a second switch in a second discharge circuit according to the output signal and the command signal by the logic circuit.

[0130] The embodiments disclosed herein are intended to be non-limiting, and those skilled in the art will recognize that specific components and arrangements of components may be modified without departing from the scope of the disclosed embodiments.

Claims

1. 1. A propulsion system for an aircraft, comprising: an electric motor configured to drive one or more propellers of the aircraft; and 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; 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, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors.

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

3. 3. A propulsion system as claimed in claim 1 or 2, wherein the first inverter circuit is controlled using midpoint reference space vector modulation.

4. A propulsion system according to any preceding claim, wherein the second inverter circuit is controlled using inverted midpoint reference space vector modulation.

5. 5. A propulsion system according to claim 1, 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. 6. The propulsion system of 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. 7. The propulsion system of claim 1, 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. 8. The propulsion system of claim 1, 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. 1. A method for controlling a propulsion system for an aircraft, comprising: stabilizing a direct current (DC) bus voltage by a capacitor; converting the DC bus voltage to an alternating current (AC) voltage according to a first pulse width modulation (PWM) vector by a first inverter circuit coupled to the capacitor to drive a first set of stator windings of an electric motor; converting the DC bus voltage to an AC voltage in response to a second PWM vector 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; driving one or more propellers of the aircraft with the electric motor.

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

11. 11. The method of claim 9 or 10, further comprising controlling the first inverter circuit using midpoint reference space vector modulation.

12. The method of any one of claims 9 to 11, further comprising controlling the second inverter circuit using inverse midpoint reference space vector modulation.

13. outputting a first set of three-phase AC voltages by the first inverter circuit to drive the first set of stator windings; 13. The method of claim 9, further comprising: outputting a second set of three-phase AC voltages by the second inverter circuit to drive the second set of stator windings.

14. 14. The method of claim 13, 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.

15. 15. The method of claim 9, further comprising reducing common mode signals on a DC side of the first inverter circuit and the second inverter circuit with a DC common mode filter coupled to the capacitor.

16. 16. The method of claim 9, further comprising reducing 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 sides of the first inverter circuit or the second inverter circuit.

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

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

19. 19. An integrated circuit according to claim 17 or 18, wherein the circuitry is further configured to control the first inverter circuit using midpoint reference space vector modulation.

20. 20. The integrated circuit of claim 17, wherein the circuitry is further configured to control the second inverter circuit using inverted midpoint reference space vector modulation.

21. the circuitry further comprising: controlling the first inverter circuit to output a first set of three-phase AC voltages by the first inverter circuit to drive the first set of stator windings; 21. The integrated circuit of claim 17, configured to control the second inverter circuit to output a second set of three-phase AC voltages by the second inverter circuit to drive the second set of stator windings.

22. 22. The integrated circuit of claim 21, 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.

23. 23. The integrated circuit of claim 17, wherein a DC common mode filter is 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.

24. 24. The integrated circuit of claim 17, wherein one or more AC common mode chokes are coupled to an AC side of the first inverter circuit or the second inverter circuit to reduce common mode signals.

25. 1. An inverter circuit for a propulsion system for an aircraft, comprising: a capacitor configured to stabilize a direct current (DC) bus voltage; a plurality of switches forming a plurality of phase legs, at least one of the phase legs including: 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; The inverter circuit, wherein the plurality of switches are controlled to short out the capacitor in response to the DC bus voltage being below a first threshold in a fault condition associated with the inverter circuit.

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

27. 27. The inverter circuit of claim 25 or 26, further comprising 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.

28. 28. The inverter circuit of claim 27, wherein the second discharge circuit comprises a discharge element and a second switch connected in series.

29. 30. The inverter circuit of claim 28, wherein the discharge element comprises a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low resistance device, or any combination thereof.

30. 30. The inverter circuit of claim 25, 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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