Inverter circuits and electric propulsion systems 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 of related applications This disclosure claims priority to U.S. Patent Application No. 18 / 363,535, filed on 1 August 2023, entitled “INVERTER CIRCUITS AND ELECTRICAL PROPULSION SYSTEMS FOR EVTOL AIRCRAFT,” which in turn claims priority to U.S. Provisional Application No. 63 / 378,536, filed on 6 October 2022, entitled “Tilt Rotor Systems and Methods for eVTOL Aircraft,” and U.S. Provisional Application No. 63 / 378,680, filed on 7 October 2022, entitled “Systems and Methods for Improved Propulsion Systems for eVTOL Aircraft.” The contents of these applications are incorporated herein by reference in their entirety for all purposes.

[0002] This disclosure generally relates to the field of powered aerial vehicles. More specifically, but not limited to, this disclosure relates to technological innovations in tiltrotor aircraft using electric propulsion systems. Certain aspects of this disclosure generally relate to improvements in electric propulsion systems for tiltrotor aircraft. Other aspects of this disclosure generally relate to improvements in power inverters, which may be used in other types of vehicles but offer particular advantages in aerial vehicles. [Overview of the project]

[0003] Embodiments of the present disclosure provide a propulsion system for an aircraft. The electric propulsion system may include: an electric motor configured to drive one or more propellers of 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 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] Embodiments of the present disclosure provide a method for controlling a propulsion system for an aircraft. The method may include: stabilizing a direct current (DC) bus voltage with a capacitor; driving the stator windings of a first set of electric motors by 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; driving the stator windings of a second set of electric motors by converting the DC bus voltage to an AC voltage in response to a second PWM vector with a second inverter circuit coupled to the capacitor, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors; and driving one or more propellers of an aircraft with the electric motors. 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 operation of the method.

[0005] Embodiments of the present disclosure provide an inverter circuit. The inverter circuit may include a capacitor configured to stabilize a direct current (DC) bus voltage and a plurality of switches forming a plurality of phase legs, at least one of the phase legs may include an upper switch positioned between the positive terminal of the capacitor and the AC output terminal of the phase leg, a lower switch positioned between the 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 under a fault condition.

[0006] Embodiments of the present disclosure provide a method for controlling an inverter circuit. The method may include detecting whether a failure has occurred in one of a plurality of switches in the inverter circuit; disconnecting the inverter circuit from the power supply in response to the detection of a single-phase short-circuit failure; providing a first discharge path by a first discharge circuit for discharging the DC bus voltage across the capacitors of the inverter circuit after the inverter circuit has been disconnected from the power supply; and controlling a plurality of switches in the inverter circuit to short-circuit the capacitors in response to the bus voltage being below a first threshold. Other embodiments may include corresponding integrated circuits, a computer system, a device, and a computer program recorded on one or more computer storage devices, each configured to perform the operation of the present method.

[0007] Embodiments of the present disclosure provide an inverter circuit. The inverter circuit may include a capacitor configured to stabilize a direct current (DC) bus voltage, and a plurality of switches forming a plurality of phase legs, wherein at least one phase leg may include an upper switch positioned between the positive terminal of the capacitor and the AC output terminal of the phase leg, a lower switch positioned between the negative terminal of the capacitor and the AC output terminal of the phase leg, a first discharge circuit coupled in parallel with 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 with 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 under a fault condition.

[0008] Embodiments of the present disclosure provide a method for controlling an inverter circuit. The method may include detecting whether a failure has occurred in one of a plurality of switches in the inverter circuit; disconnecting the inverter circuit from the power supply in response to the detection of a single-phase short-circuit fault; and after the inverter circuit has been disconnected from the power supply, discharging the bus voltage across the capacitors of the inverter circuit by providing a first discharge path using a first discharge circuit in response to confirmation that the inverter circuit has been disconnected from the power supply, and by providing a second discharge path in parallel with the first discharge path using a second discharge circuit in response to the bus voltage being below a threshold. Other embodiments may include corresponding integrated circuits, a computer system, an apparatus, and a computer program recorded on one or more computer storage devices, each configured to perform the operation of the method. [Brief explanation of the drawing]

[0009] [Figure 1] An illustrative perspective view of an exemplary VTOL aircraft, consistent with some embodiments of the present disclosure. [Figure 2] Another illustrative perspective view of an exemplary VTOL aircraft with an alternative configuration consistent with some embodiments of this disclosure. [Figure 3] This is an illustrative top view of an exemplary VTOL aircraft, consistent with some embodiments of the present disclosure. [Figure 4] This schematic diagram illustrates exemplary propeller rotation of a VTOL aircraft, consistent with some embodiments of the present disclosure. [Figure 5] This schematic diagram illustrates an exemplary power connection in a VTOL aircraft, consistent with some embodiments of the present disclosure. [Figure 6] This is a block diagram illustrating an exemplary architecture and design of an electric propulsion unit for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 7] This is a schematic diagram illustrating an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 8A] This is an illustrative example of an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 8B] This is an illustrative example of an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 8C] This is an illustrative example of an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 9] This is a schematic diagram illustrating an exemplary electric ascent propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 10A] This is an illustrative example of an exemplary electric ascent propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 10B] This is an illustrative example of an exemplary electric ascent propulsion system for a VTOL aircraft, consistent with embodiments of the present disclosure. [Figure 11] This figure 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] Figure 11 illustrates a PWM vector for controlling an inverter circuit in an electric propulsion system, consistent with some embodiments of the present disclosure. [Figure 12B]A diagram illustrating three-phase voltages output by an inverter circuit in one cycle, which is consistent with some embodiments of the present disclosure. [Figure 13A] A diagram illustrating an inverter circuit for motor control in an electric propulsion system, which is consistent with some embodiments of the present disclosure. [Figure 13B] A diagram illustrating another inverter circuit for motor control in an electric propulsion system, which is consistent with some embodiments of the present disclosure. [Figure 14] A diagram illustrating an exemplary flowchart of a method for controlling the inverter circuit of FIG. 13A or FIG. 13B, which is consistent with some embodiments of the present disclosure. [Figure 15] A graph showing the bus voltage over time for the HV DC bus during the discharge period, which is consistent with some embodiments of the present disclosure. [Figure 16A] A graph showing EMI noise in a dual-inverter system or a single-inverter system, which is consistent with some embodiments of the present disclosure. [Figure 16B] A graph showing EMI noise in a dual-inverter system or a single-inverter system, which is consistent with some embodiments of the present disclosure. [Figure 16C] A graph showing EMI noise in a dual-inverter system or a single-inverter system, which is consistent with some embodiments of the present disclosure. [Figure 16D] A graph showing EMI noise in a dual-inverter system or a single-inverter system, which is consistent with some 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. For the sake of simplifying the present disclosure, specific examples of components and arrangements are described below. These are, of course, merely examples and are not intended to be limiting. Additionally, the present disclosure may repeat reference numerals and / or letters in various examples. This repetition is for the purpose of simplification and clarity and does not in itself determine the relationship between the various embodiments and / or configurations being 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 in this specification, including examples of any terms discussed herein, is for illustrative purposes only and in no way limits the scope and meaning of the present disclosure or any of the exemplary terms. Similarly, the present disclosure is not limited to the various embodiments provided herein.

[0012] Terms such as "first," "second," etc. 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 may be referred to as a second element, and similarly, a second element may 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] Furthermore, spatially relative terms, such as “below,” “below,” “on the lower side,” “above,” and “on the upper side,” may be used herein to facilitate descriptions of the relationship between one element or feature illustrated in the figures and another element or feature. Spatially relative terms are intended to encompass different orientations of the device in use or operation, in addition to the orientation shown in the figures. The device may be in other orientations (rotated 90 degrees or otherwise), and the spatially relative descriptors used herein may be interpreted accordingly.

[0014] In this book, the term “combined” may also be referred to as “electrically coupled,” and the term “connected” may also be referred to as “electrically connected.” “Combined” and “connected” may also be used to indicate that two or more elements cooperate or interact with each other.

[0015] This disclosure addresses components of electric vertical take-off and landing (eVTOL) aircraft, primarily for use in non-conventional aircraft. For example, the eVTOL aircraft of this 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 densely populated areas. The aircraft may be intended to carry 4 to 6 passengers or commuters who expect a low-noise and low-vibration experience. Therefore, it may be desirable that the aircraft components be configured and designed to withstand frequent use without wear, that the components generate less heat and vibration, and that the aircraft 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 that their components be configured and designed to generate low levels of noise 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 able to perform conventional takeoffs and landings on runways. Furthermore, it may be desirable for an aircraft to be able to safely take off and land vertically from relatively confined spaces (e.g., vertiports, parking lots, or private roads) while transporting approximately 4-6 passengers or commuters with accompanying baggage, compared to conventional airport runways. These usage requirements may impose design constraints on the aircraft's size, weight, and operational efficiency (e.g., drag, energy use), which can affect the design and configuration of the aircraft's components.

[0016] The disclosed embodiments provide novel and improved configurations of aircraft components not observed in conventional aircraft, and / or identified design criteria for components that differ from those of conventional aircraft. Such alternative configurations and design criteria, in combination, address the shortcomings and challenges of conventional components, giving rise to the embodiments disclosed herein for various configurations and designs of eVTOL aircraft components.

[0017] In some embodiments, the eVTOL aircraft of this disclosure may be designed to be capable of both vertical and conventional takeoff and landing, having a distributed electric propulsion system that enables vertical flight, forward flight, and transition. Thrust may be generated by supplying high-voltage power to electric engines of the distributed electric propulsion system, each of which may convert the high-voltage power into mechanical shaft power to rotate a propeller. Embodiments disclosed herein may involve optimizing the energy density of the electric propulsion system. Embodiments may include electric engines connected to an on-board 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 utilizing or generating electricity, such as a fuel-powered generator or a solar panel array. Some disclosed embodiments provide weight reduction and space reduction of components in the aircraft, thereby increasing the efficiency and performance of the aircraft. Focusing on safety in passenger transport, the disclosed embodiments also implement new and improved safety protocols and system redundancy in the event of failure to minimize any single point of failure in the aircraft propulsion system. Some of the disclosed embodiments also provide new and improved approaches to meeting aviation and transport laws and regulations. For example, the Federal Aviation Administration enforces federal laws and regulations that require safety components, such as fire barriers, adjacent to engines that use quantities of oil or other flammable materials exceeding a threshold.

[0018] In a preferred embodiment, the distributed electric propulsion system may include twelve electric engines that can be mounted on forward and aft booms of the aircraft's wings. The forward electric engines may be tiltable during flight between a horizontal position (e.g., for generating forward thrust) and a vertical position (e.g., for generating 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 vertical position (e.g., for generating 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 are not equal. In some embodiments, the aircraft may have four forward propellers and four rear propellers, at least four of which include tiltable propellers.

[0019] In a preferred embodiment, for vertical takeoff and landing (VTOL) missions, the forward and rear electric engines may provide vertical thrust during takeoff and landing. During the flight phase when the aircraft is in forward flight mode, the forward electric engines may provide horizontal thrust, while the rear electric engine propellers may be retracted to a fixed position to minimize drag. The rear electric engines may be actively retracted while maintaining position monitoring. Transitions from vertical to horizontal flight and vice versa may be achieved via a tilt propeller subsystem. The tilt propeller subsystem may redirect thrust, primarily vertical, during vertical flight mode to a nearly horizontal direction during forward flight. A variable pitch mechanism may change the collective angle of the propeller hub assembly blades of the forward electric engine for operation during the hovering, transition, and cruising phases.

[0020] In some embodiments, in conventional take-off and landing (CTOL) missions, the forward electric engines may provide horizontal thrust for fixed-wing take-off, cruising, and landing. In some embodiments, the rear electric engines may not be used to generate thrust during CTOL missions, and the rear propellers may be retracted into a fixed position.

[0021] In some embodiments, the electric engine may be housed in or connected to the boom of an aircraft and may include a motor, an inverter, and a gearbox. In some embodiments, the motor, inverter, and gearbox may be linked so as to share a central axis. In some embodiments, the torque generated by the motor may be sent to the gearbox, separate from the propeller of the propulsion system. In some embodiments, the gearbox may provide gear reduction and then send the torque back to the propeller via the main shaft and bearings located inside the motor. In some embodiments, the inverter may be mounted on the back of the gearbox so as to prevent the main shaft from advancing through the inverter when outputting torque to the propeller. In some embodiments, the motor, gearbox, and inverter may be linked so as 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 including an amount less than one quart, two quarts, three quarts, or any other metered amount of oil.

[0022] In some embodiments, the tilt propeller system may include a linear actuator or a 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 work together to provide gear reduction that can orient the propulsion system. In some embodiments, the tilt propeller system may include a redundant configuration in which multiple motors, inverters, and gearboxes are present and work together using gears. In some embodiments, a configuration utilizing multiple motors, gearboxes, and inverters may allow a failed part of the redundant configuration to be driven by the motors, inverters, and gearboxes of another part of the configuration. In some embodiments, a gearbox configuration may also allow the tilt 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) power to an electric engine, which converts the HV power into mechanical shaft power used to rotate a propeller. As described above, the aircraft described herein may have a plurality of boom-mounted electric engines on the front and rear of the wings. 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 the orientation or tilt of the forward electric engine may be changeable. Additional embodiments include a forward engine, which may be of 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 a rear engine or lifter, which may be of the clockwise (CW) or counterclockwise (CCW) type. Additional embodiments may include a rear electric engine utilizing a multi-blade fixed-pitch propeller.

[0025] As described herein, the orientation and use of the electric propulsion system may vary throughout the aircraft's operation. In some embodiments, during vertical takeoff and landing, the forward propulsion system, and the rearward propulsion system, may provide vertical thrust during takeoff and landing. During the flight phase when the aircraft is in forward flight mode, the forward propulsion system may provide horizontal thrust, while the propellers of the rearward propulsion system may be retracted to a fixed position to minimize drag. The rearward electric propulsion system may be actively retracted while monitoring its position. Some embodiments may include transitions from vertical to horizontal flight and vice versa. In some embodiments, the transition may be achieved via a tilt propeller system (TPS). The TPS redirects thrust, primarily vertical, during vertical flight mode, primarily horizontal, during forward flight mode. Additional embodiments may include a variable pitch mechanism that can change the collective angle of the propeller hub assembly blades of the forward propulsion system for operation during the hovering phase, cruising phase, and transition phase. Some embodiments may include a conventional takeoff and landing (CTOL) configuration in which the tilter provides horizontal thrust for fixed-wing takeoff, cruising, and landing. The rear electric engine is not used to generate thrust during CTOL missions, and the rear propeller is retracted into its fixed position.

[0026] In some embodiments, the electric engines described herein may have design features to mitigate and protect against uncontained fires, such as not having a nominal ignition source within the electric engine and utilizing a non-hazardous amount of flammable fluid contained in both tilt and lift engines, which may have an engine heating operating limit that is more than 50°C lower than the self-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 allow the electric engine to be considered not to be a designated fire zone. In some embodiments, the flammable fluid may include oil, and the non-hazardous amount may be less than 1 quart, 2 quarts, 3 quarts, 4 quarts, 5 quarts, or 10 quarts, depending 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 motor, or an inverter, gearbox, and motor, in a variety of configurations, such as the typical configurations described herein. For example, an electric engine may include an electric motor, gearbox, and inverter, all sharing the same central axis. In addition, the central axis may be configured along the axis of an output shaft directed toward an aircraft propeller. In such exemplary configurations, the motor, gearbox, and inverter would all share the output shaft as the central axis and would be oriented circularly around the output shaft. Additional embodiments may include a motor, gearbox, and inverter, which are mounted together in a certain arrangement, or some of the components, such as the motor and gearbox, are mounted together, while other components, such as the inverter, are located elsewhere, but are mounted together in a configuration in which a wiring system is used to connect the electric engine.

[0028] As described above, the electric engines for aircraft described herein may include some or all of a motor, inverter, and gearbox. Various configurations may include an inverter and motor such that the motor's output shaft directly provides speed and torque to the propeller shaft. Additional embodiments of the electric engine may include a motor, inverter, and gearbox in which the motor's output may be transmitted through a gearbox connected to an output shaft to the propeller, or a motor, inverter, and gearbox in which the output from the motor is transmitted through the gearbox, decoupled from the propeller, and the output shaft to the propeller then proceeds to the propeller through the gearbox and motor. As described herein, the electric engine may consider any combination or orientation of some or all of the motor, inverter, and gearbox. In addition, each configuration or orientation of the electric engine disclosed herein may include cooling via air cooling, coolant, or a mixture of both.

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

[0030] Some embodiments may include an electric engine in which the inverter module may be mounted outside the motor enclosure. Additional embodiments may include an electric engine in which the inverter may be mounted above the electric motor such that the cooling fins of the inverter are below the propeller. Further embodiments may include an inverter mounted on the rear of the motor such that the cooling fins face radially outward, an inverter mounted on the front of the motor such that the cooling fins face radially outward, and an inverter mounted on the motor in which the inverter is cooled by a liquid such as oil, or by any other location of the inverter relative to the motor.

[0031] Embodiments of an electric motor may include a stator enclosure, a wound stator assembly, a rotor, various bearings, and any additional components that help transmit 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 be equipped with a thermal management system to ensure that the components of the electric engine do not fail during operation. In some embodiments, a coolant may be used and circulated throughout the individual components of the engine, such as an inverter, gearbox, or motor, through some of the components, or through all of the components of the engine, to help manage the heat present in the engine. Additional embodiments may include using an air cooling method to cool the electric engine, or using a mixture of coolant and air to manage the heat generated in 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 in the inverter and gearbox, or any other combination of air and liquid cooling over the inverter, gearbox, and motor, or even a subset of those components.

[0033] In some embodiments, oil may be used as a lubricant throughout the electric engine and as a coolant fluid to help manage the heat generated by the engine during operation. In addition to this example, various amounts of oil, such as less than one quart, less than two quarts, or any other amount necessary to lubricate and cool the electric engine, may be used in combination with or without air cooling assistance to function as both a lubricant and a coolant fluid in an electric engine. As disclosed herein, electric engines may have different primary functions, such as being used only for climbing and landing and thus only in one orientation, or being used in all stages of flight, such as climbing, landing, and during flight. An engine used in all stages of flight may experience various orientations throughout the flight and may contain more lubricant and coolant than an engine used in only one orientation. Thus, not all engines on an aircraft have to contain the same amounts of lubricant and coolant. For example, a climb and landing engine may require less than one quart of oil, while an engine operating in all stages 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 define limits on the amounts of lubricants and coolants that may be used in electric engines.

[0034] While using oil to cool an electric engine, rather than a separate coolant, adds an additional oil to the system, it should be understood that the oil eliminates the need for conventional components that could be used to cool such an electric engine. For example, if an electric engine is cooled by another liquid such as glycol, the engine may have separate heat exchangers for both the lubricating fluid and the coolant fluid. Thus, in embodiments where a single fluid such as oil is used for both lubrication and cooling, there will be an increase in oil, but only the need for one heat exchanger will exist, and therefore, in the overall system, there may be a reduction in mass due to fewer heat exchangers used and other potentially unnecessary components, and a more attractive drag profile may exist. Furthermore, using one substance for engine lubrication and cooling can increase the efficiency of the system due to the reduction in mass, as well as the advantages of cooling the engine with a substance rather than relying on air cooling, which can be problematic throughout the engine.

[0035] Additional embodiments of electric engines may have various components to ensure that any flammable fluid is monitored and prevented from entering specific sections of the electric engine. Some embodiments may include an electric engine having a wet zone enclosure which may be defined by a gearbox, motor, and / or heat exchanger. In some embodiments, an 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, an electric engine may have up to 5 liters, or 6 liters, or 8 liters, or 10 liters, or 20 liters of air in the motor gearbox enclosure in contact with the engine oil, based on factors such as the size of the aircraft, the number of propellers, or the payload. Embodiments of the motor gearbox enclosure may use a breather to equalize the internal and external pressures. Embodiments of the breather may include the breather protruding over nearby design features to prevent accidental ingress of external fluids. Additional embodiments may include a breather with a screen and a bypass ingress path to prevent ingress of external debris. Embodiments may include sight glasses present in both the tilt electric engine and the rise electric engine to ensure that the oil is not overfilled or underfilled during maintenance.

[0036] Additional embodiments of the electric engine may include active protection features in forward and rear electric engines, such as monitoring internal temperatures throughout the engine, including oil temperature, stator winding set, inverter bulk capacitor, power module, control board power module, control board control processor, control board monitor processor, 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 the self-ignition temperature of the fluid. Some embodiments may include a high-voltage power system that may have fuses at high-voltage battery terminals that can quickly disconnect the engine electrical connections to mitigate irreversible overcurrent events. This overcurrent protection may be activated when the current consumption of the electric engine is greater than the overcurrent operation. Thus, in some embodiments, fault conditions that would lead to overcurrent may only lead to transient overheating, arcing, or sparking faults. Some embodiments may include a fire threat characterization test ignition source that can 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 detects AC overcurrents, isolates misphases, and / or continuously monitors the input DC voltage and applies protective actions to maintain the voltage below the overvoltage operating limit.

[0037] A. Characteristics of an exemplary electric aircraft Figure 1 is an illustrative perspective view of an exemplary VTOL aircraft consistent with the disclosed embodiments. Figure 2 is another illustrative perspective view of an exemplary VTOL aircraft in an alternative configuration consistent with embodiments of the present disclosure. Figures 1 and 2 show VTOL aircraft 100, 200 in cruising configuration and vertical takeoff, landing, and hovering configuration (also referred to herein as “climb” configuration), respectively, consistent with embodiments of the present disclosure. Elements corresponding to Figures 1 and 2 may have similar figures and may refer to similar elements of aircraft 100, 200. Aircraft 100, 200 may include fuselages 102, 202, wings 104, 204 mounted on fuselages 102, 202, and one or more rear stabilizers 106, 206 mounted on the rear of fuselages 102, 202. Multiple lift propellers 112, 212 may be mounted on wings 104, 204 and configured to provide lift for vertical takeoff, landing, and hovering. Multiple tilt propellers 114, 214 may be mounted on wings 104, 204 and may be tiltable between an ascent configuration, as shown in Figure 2, which provides some of the lift required for vertical takeoff, landing, and hovering, and a cruising configuration, as shown in Figure 1, which provides forward thrust to the aircraft 100 for horizontal flight. As used herein, the ascent configuration of a tilt propeller refers to any tilt propeller orientation in which the tilt propeller thrust primarily provides lift to the aircraft, and the cruising configuration of a tilt propeller refers to any tilt propeller orientation in which the tilt propeller thrust primarily provides forward thrust to the aircraft.

[0038] In some embodiments, the lift propellers 112, 212 may be configured to provide only lift, with all horizontal thrust 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. On the other hand, the tilt propellers 114, 214 may be tilted upward in an upward configuration, where the thrust from the propellers 114, 214 is directed downward to provide additional lift.

[0039] For forward flight, the tilt propellers 114, 214 can be tilted from their ascent configuration to their cruising configuration. In other words, the orientation of the tilt propellers 114, 214 can change from an orientation in which the thrust of the tilt propellers is directed downward (to provide lift during vertical takeoff, landing, and hovering) to an orientation in which the thrust of the tilt propellers is directed aft (to provide forward thrust to the aircraft 100, 200). The tilt propeller assembly for a particular electric engine can be tilted around an axis of rotation defined by the mounting point connecting the boom and the electric engine. When the aircraft 100, 200 is in full forward flight, lift can be fully provided by the wings 104, 204. On the other hand, in the cruising configuration, the lift propellers 112, 212 can be shut off. The blades 120, 220 of the lift propellers 112, 212 can be held in a low-drag position for aircraft cruising. 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 where one blade is directly in front of the other blade, as illustrated in Figure 1. In some embodiments, the lift propellers 112, 212 may 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 Figures 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, each of the tilt propellers 114, 214 may have two to five blades, and in some cases more blades, 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 over the entire aircraft). At least some of the lift propellers 112, 212 may be located behind the wings 104, 204, and at least some of the tilt propellers 114, 214 may be located in front of the wings 104, 204. In some embodiments, all of the lift propellers 112, 212 may be located behind the wings 104, 204, and all of the tilt propellers 114, 214 may be located in front of the wings 104, 204. According to some embodiments, all of the lift propellers 112, 212 and tilt propellers 114, 214 may be mounted on the wings, i.e., the lift propellers or tilt propellers may not be mounted on the fuselage. In some embodiments, all lift propellers 112, 212 may be located behind the wings 104, 204, and all tilt propellers 114, 214 may be located in front of the wings 104, 204. According to some embodiments, all lift propellers 112, 212 and tilt propellers 114, 214 may be located inside the ends of the wings 104, 204.

[0041] In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted on the wings 104, 204 by booms 122, 222. Booms 122, 222 may be mounted below, above, and / or incorporated into the wing profile. In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted directly on the wings 104, 204. In some embodiments, one lift propeller 112, 212 and one tilt propeller 114, 214 may be mounted on each boom 122, 222. Lift propellers 112, 212 may be mounted at the rear end of booms 122, 222, and tilt propellers 114, 214 may be mounted at the front end of booms 122, 222. In some embodiments, lift propellers 112, 212 may be mounted in fixed positions on booms 122, 222. In some embodiments, tilt propellers 114, 214 may be mounted to the front end of booms 122, 222 via hinges. The tilt propellers 114, 214 may be mounted on booms 122, 222 such that when in their cruising configuration, the tilt propellers 114, 214 are aligned with the body of booms 122, 222 and form a continuous extension of the front end of booms 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 wings 104, 204 may be designed with a profile that reduces drag during forward flight. In some embodiments, the wingtip profile may be curved and / or tapered to minimize drag.

[0043] In some embodiments, the rear stabilizers 106, 206 include control surfaces such as one or more rudders, one or more elevators, and / or one or more combined rudder-elevators. The wings(s) may have any suitable design. In some embodiments, the wings have tapered leading edges.

[0044] In some embodiments, a lift propeller 112, 212 or tilt propeller 114, 214 may 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 axis of rotation of a lift propeller / tilt propeller around a line parallel to the longitudinal direction, similar to the roll degrees of freedom of an aircraft. The tilting of a lift propeller and / or tilt propeller may help minimize damage from propeller rupture by oriented the plane of rotation of the lift propeller / tilt propeller disc (the blades and the hub on which those blades are mounted) so as not to intersect critical parts of the aircraft (such as areas of the fuselage where a person may be positioned, critical flight control systems, batteries, adjacent propellers, etc.) or other propeller discs, and may provide enhanced yaw control during flight.

[0045] Figure 3 is an illustrative 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 aircraft 100 and 200 shown in Figures 1 and 2, respectively. As discussed herein, the aircraft 300 may include twelve electric propulsion systems distributed across the aircraft 300. In some embodiments, the distribution of electric propulsion systems may include six forward electric propulsion systems 314 and six rear electric propulsion systems 312 mounted on the forward and rear booms of the main wing 304 of the aircraft 300. In some embodiments, the length of the trailing end of the boom 324 from the wing 304 to the lift propeller may include similar trailing ends of the boom 324 across a number of trailing ends of the boom. In some embodiments, the length of the trailing end of the boom may vary across six exemplary trailing ends of the boom. For example, each trailing end of the boom 324 may have a different length from the wing 304 to the lift propeller, or a subset of the trailing ends of the boom may have similar lengths. In some embodiments, the front end of the boom 322 may include varying lengths from the wing 304 to the tilt propeller across the front end of the boom. For example, as shown in Figure 3, the length of the front end of the boom 322 from the tilt propeller closest to the fuselage to the wing 304 may include a longer length from the wing 304 to the tilt propeller furthest from the fuselage. Some embodiments may include boom fronts having similar lengths across six exemplary front ends of the boom, or any other distribution of boom front lengths from the wing 304 to the tilt propeller. Some embodiments may include an aircraft 300 having eight electric propulsion systems, including four forward electric propulsion systems 314 and four rear electric propulsion systems 312, or any other distribution of forward and rear electric propulsion systems, including embodiments in which the number of forward electric propulsion systems 314 is less than or greater than the number of rear electric propulsion systems 312. Furthermore, Figure 3 shows an exemplary embodiment of a VTOL aircraft 300, which has forward propellers oriented horizontally for horizontal flight and rear propeller blades 320 in a retracted position for forward flight.

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

[0047] Figure 4 is a schematic diagram illustrating exemplary propeller rotation of a VTOL aircraft, consistent with the disclosed embodiments. The aircraft 400 shown in the figure may be a top view of aircraft 100, 200, and 300 shown in Figures 1, 2, and 3, respectively. The aircraft 400 may include six forward electric propulsion systems, three of which are CW type 424 and the remaining three forward electric propulsion systems are CCW type. In some embodiments, three rearward electric propulsion systems may be CCW type 428 and the remaining three rearward electric propulsion systems are CW type 430. Some embodiments may include an aircraft 400 having four forward electric propulsion systems and four rearward electric propulsion systems, each having two CW type and two CCW type. In some embodiments, propellers may be reversed relative to adjacent propellers to cancel torque steer generated by the propeller rotation and experienced by the aircraft's fuselage or wings. In some embodiments, the difference in rotation direction may be achieved using the 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 having a forward electric propulsion system and a rear electric propulsion system, where the quantities of CW type 424 and CCW type 426 are not equal between forward electric propulsion systems, between rear electric propulsion systems, or between forward electric propulsion systems and rear electric propulsion systems.

[0049] Figure 5 is a schematic diagram illustrating exemplary power connections in a VTOL aircraft, consistent with the disclosed embodiments. A 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. In some embodiments, high-voltage power systems may be 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 the 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 rearward electric propulsion systems having six electric engines 514, 516, 518, 520, 522, and 524. In some embodiments, batteries may be connected to diagonally opposed electric engines. In such a configuration, the first power system 526 may supply power to the electric engine 502 via a power connection channel 538 and to the electric engine 524 via a power connection channel 540. In some embodiments, the first power system 526 may be paired with a fourth power system 532 via a power connection channel 542 having a fuse to prevent excessive current from flowing through power systems 526 and 532. In addition to this embodiment, the VTOL aircraft 500 may include a second power system 528 paired with a fifth power system 534 via a power connection channel 548 having a fuse, which may supply power to electric engines 510 and 516 via power connection channels 544 and 546, respectively. In some embodiments, a third power system 530 may be paired with a sixth power system 536 via a power connection channel 554 having a fuse, and may supply power to electric engines 506 and 520 via power connection channels 550 and 552, respectively. A fourth power system 532 may also supply power to electric engines 508 and 518 via power connection channels 556 and 558, respectively. A fifth power system 534 may also supply power to electric engines 504 and 522 via power connection channels 560 and 562, respectively.The sixth power grid 536 can also supply power to the 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 a battery located within the aircraft, via a high-voltage channel or power connection channel. Some embodiments may include various batteries housed within the aircraft wing, having high-voltage channels that run to the electric propulsion system throughout the entire 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, the aircraft may include multiple electric propulsion systems that can be pattern-wired to various batteries or power sources housed throughout the aircraft. It is recognized that such a configuration may be beneficial in avoiding the risk of a single point of failure, where the failure of one battery or power source could result in a portion of the aircraft being unable to maintain the amount of thrust necessary to continue flight or perform a controlled landing. For example, if a VTOL has two forward electric propulsion systems and two rearward electric propulsion systems, the forward and rearward electric propulsion systems 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 rear electric propulsion systems on the opposite side of the VTOL aircraft may remain operational, providing a more balanced flight or landing compared to a failed forward and rear electric propulsion system on the same side of the VTOL aircraft. Some embodiments may include four forward and four rear electric propulsion systems in which diagonally opposed electric engines are connected to a common battery or power source. Some embodiments may include various configurations of electric engines electrically connected to a high-voltage power system so that the risk of a single point of failure in the event of a power failure can be avoided, and the flight phase in which the failure occurs can continue, or the aircraft can perform an alternative flight phase 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 the electric propulsion system may include a high-voltage power system that supplies high-voltage power to the electric engine and / or a low-voltage system that supplies low-voltage DC power to the electric engine. Some embodiments may include an electric engine(s) that digitally communicates with a flight control system ("FCS") which includes a flight control computer ("FCC") that can send and receive signals to and from the electric engine, including command and response data or status. Some embodiments may include an electric engine that can receive operating parameters from the FCC, including speed, voltage, current, torque, temperature, vibration, propeller position, and any other values ​​of operating parameters, and transmit the operating parameters to the FCC.

[0052] In some embodiments, the flight control system may include a system capable of communicating with the electric engine to send and receive analog / discrete signals to the electric engine and controlling a device that can redirect the thrust of the tilt propeller between primarily vertical in vertical flight mode and primarily horizontal in forward flight mode. In some embodiments, this system may be referred to as a tilt propeller system ("TPS") and may be capable of transmitting and oriented additional features of the electric propulsion system.

[0053] Figure 6 illustrates an exemplary architecture and design block diagram of an electric propulsion unit 600 consistent with the disclosed embodiments. In some embodiments, the electric propulsion system 602 may include an electric engine subsystem 604 capable of supplying torque to a propeller subsystem 606 via a shaft to generate thrust for the electric propulsion system 602. In some embodiments, the electric engine subsystem 604 may receive low-voltage DC (LV DC) power from a low-voltage system (LVS) 608. In some embodiments, the electric engine subsystem 604 may receive high-voltage (HV) power from a high-voltage power system (HVPS) 610, which 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 supplying high-voltage power to the electric engine subsystem 604. It is recognized that such configurations may be advantageous in that they do not jeopardize a single point of failure where the failure of a single battery would lead to the failure of the electric propulsion system 602.

[0054] In some embodiments, the electric propulsion system 602 may include an electric engine subsystem 604 that receives signals from and transmits signals to the 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 the CAN data bus signals are used between the flight control computer and the electric engine(s), it should be understood that in some embodiments, any form of communication capable of sending and receiving data from the flight control computer to and from the electric engine(s) may be included. 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 a device that transmits operating parameters to the electric engine subsystem 604 and articulates the orientation of the propeller subsystem 606, thereby redirecting the thrust of the tilt propeller during various phases of flight using mechanical means such as a gearbox assembly, a linear actuator, and any other configuration of components for changing the orientation of the propeller subsystem 606.

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

[0056] Figure 7 is a schematic diagram illustrating an exemplary tilt electric propulsion system for a VTOL aircraft, consistent with the 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 having a hub, spinner, and tilt propeller blades. In some embodiments, the electric engine assembly 702 may include a motor and gearbox assembly 704 aligned along the shaft 724 and mechanically coupled. In some embodiments, the motor and gearbox assembly 704 may include an electric motor assembly having a stator 706 and a rotor 708. As shown in Figure 7 and in some embodiments, the stator 706 may include a plurality of stator windings connected to an inverter 716. In such a configuration, the stator 706 may incorporate one or more redundancies, so that if one or more sets of windings fail, power is still transmitted to the stator 706 through one or more remaining windings, and as a result, the electric engine assembly 702 retains power and continues to generate thrust in the propeller assembly 720.

[0057] In some embodiments, the motor and gearbox assembly 704 may include a gearbox 710 aligned along the shaft 724 to provide gear reduction between the torque on the shaft 724 from the electric engine assembly, which comprises a stator 706 and a rotor 708, and the output shaft 738. The torque applied to the output shaft 738 can be transmitted to the propeller assembly 720. In some embodiments, the gearbox 710 may include an oil pump. In such embodiments, the oil pump may drive the circulation of oil throughout the motor and gearbox assembly 704 at a speed equal to the rotation of the output shaft 738 to cool and lubricate the gearbox and electric motor components. In some embodiments, the oil pump may drive the circulation of oil at a speed greater or less than the rotation of the output shaft 738. In some embodiments of the motor and gearbox assembly 704, a propeller position sensor 712 located within the housing may include a propeller position sensor 712 that can detect the magnetic field generated by the electric engine assembly to determine the 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, the electric engine assembly 702 may also include an inverter assembly 714 substantially aligned along the shaft 724. The inverter assembly 714 may include an inverter 716 and an inverter power supply 740. The inverter power supply 740 may accept low-voltage DC power from a low-voltage system 734 located outside the electric engine assembly 702. The inverter power supply 740 may accept low-voltage DC power originating from a high-voltage power system 732 located outside the electric engine assembly 702, which has been converted to low-voltage DC power via a DC-DC converter 742. The inverter 716 may supply high-voltage alternating current (AC) to the stator 706 of the electric engine assembly located within the motor and gearbox assembly 704 via at least one three-phase winding. The inverter assembly 714 may include an inverter 716 capable of receiving flight control data from a flight control computing subsystem 736.

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

[0060] In some embodiments, the electric engine assembly 702 may be mounted to or coupled to the aircraft's boom structure 726. The variable pitch mechanism 730 may be mechanically coupled to the propeller assembly 720. In some embodiments, the variable pitch mechanism may abut against the electric engine assembly 702. In some embodiments, the variable pitch mechanism 730 may be coupled to the variable pitch mechanism 730 so that it can be remotely mounted in the aircraft's boom, wing, or fuselage. In some embodiments, the variable pitch mechanism 730 may include a shaft or component that proceeds into the propeller assembly 720 within or adjacent to the shaft 724. The variable pitch mechanism 730 may help to change the collective angle of the forward electric engine's propeller hub assembly blades as needed for operation during the hovering, transition, and cruising phases. In some embodiments, the electric engine assembly 702 may include being mechanically coupled to a tilt propeller subsystem 728 that can redirect thrust between primarily vertical during vertical flight mode and primarily horizontal during forward flight mode. In some embodiments, the tilt propeller subsystem may be in contact with the variable pitch mechanism 730. Some embodiments may include a tilt propeller subsystem 728 comprising various components located in different locations. For example, components of the tilt propeller subsystem may be coupled to the electric engine assembly 702, while 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] Figures 8A–8C illustrate exemplary tilt electric propulsion systems for VTOL aircraft, consistent with the disclosed embodiments. Figures 8A–8C refer to similar elements of tiltable electric propulsion systems 800A, 800B, and 800C, which have similar numbering. Thus, similar design considerations and configurations may be considered throughout the embodiments.

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

[0063] Figure 8C illustrates a top-down view along the spinner 808C of a tiltable electric propulsion system 800C in an ascent configuration integrated into boom 812B, consistent with the present disclosure. As shown in Figure 8C, the tiltable electric propulsion system 800C in the ascent configuration may include electric engine assemblies 802A, 802B that are located outside the boom 812C and whose orientation relative to the boom 812C is altered.

[0064] As discussed herein, an electric ascent 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 may not provide thrust during cruising.

[0065] Figure 9 is a schematic diagram illustrating an exemplary climb electric propulsion system for a VTOL aircraft, consistent with the disclosed embodiments. The climb electric propulsion system 900 may be mounted on or coupled to the boom structure 924 of the aircraft. The climb 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 having a hub and tilt propeller blades. In some embodiments, the electric engine assembly 902 may include a motor and gearbox assembly housing 904 aligned along the shaft 940 and mechanically coupled. In some embodiments, the motor and gearbox assembly housing 904 may include an electric motor assembly having a stator 906 and a rotor 908. The stator 906 may include a plurality of stator windings connected to an inverter 916. In such configurations, the stator 906 may incorporate one or more redundancies and backup measures to avoid a single point of failure in the case. For example, the stator 906 may include multiple windings so that if one winding fails, power can continue to be transmitted to the stator 906 through the remaining windings, allowing the electric engine assembly 902 to retain power and continue generating thrust in the propeller assembly 920.

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

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

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

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

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

[0071] Figures 10A and 10B are illustrative examples of an exemplary climb electric propulsion system for a VTOL aircraft, consistent with the disclosed embodiments. Figures 10A and 10B share similar figures and refer to similar elements of climb electric propulsion systems 1000A and 1000B. Thus, similar design considerations and configurations may be considered throughout the embodiments.

[0072] Figure 10A illustrates a side profile of a climb electric propulsion system 1000A in a climb configuration integrated into boom 1010A, consistent with the present disclosure. The climb electric propulsion system 1000A may comprise an electric engine assembly 1002A housed within boom 1010A of a VTOL aircraft. In some embodiments, the climb configuration may include an electric engine assembly 1002A positioned vertically within 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] Figure 10B illustrates a top-down view of the lift electric propulsion system 1000B in a lift configuration integrated into the boom 1010B, consistent with the present disclosure.

[0074] Some embodiments of the disclosed electric engine may generate heat during operation and may be equipped with a thermal management system to ensure that the components of the electric engine do not fail during operation. In some embodiments, a coolant may be used and circulated throughout the individual components of the engine, such as an inverter, gearbox, or motor, through some of the components, or through all of the components of the engine, to help manage the heat present in the engine. Some embodiments may include using an air cooling method to cool the electric engine, or using a mixture of coolant and air to manage the heat generated in 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, 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 the cooling mixture may be used in 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 the heat generated by the engine during operation. In addition to this example, different amounts of oil may be used to function as both a lubricant and a coolant fluid in the electric engine, in combination with any other amount of oil necessary to lubricate and cool the electric engine, such as less than 1 quart, 1.5 quarts, 2 quarts, 2.5 quarts, 3 quarts, 5 quarts, or with or without air cooling support. 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 required 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 climbing and landing and thus only in one orientation, or being used in all stages of flight, such as climbing, landing, and during flight. An engine used in all stages of flight may experience various orientations throughout the flight and may contain more lubricant and coolant than an engine used in only one orientation. Thus, not all engines on an aircraft have to contain the same amount of lubricant and coolant. For example, climb and landing engines may require less than one quart of oil, while engines operating at all stages of flight may require more than one quart. In some embodiments, the amount of cooling oil or liquid may be appropriate to provide enough heat 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 illustrative and non-limiting and do not determine limits on the amounts of lubricants and coolants that may be used in electric engines.

[0076] In some embodiments, oil may be used 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 could be used to cool such an electric engine. For example, if the electric engine is cooled by another liquid such as glycol, the engine may have separate heat exchangers for both the lubricating fluid and the coolant fluid. Thus, in embodiments where a single fluid such as oil is used for both lubrication and cooling, there will be an increase in oil, but only the need for one heat exchanger will exist, and therefore, in the overall system, there may be a reduction in mass due to fewer heat exchangers used and other components that are potentially unnecessary, and a more attractive drag profile may exist. Furthermore, using one substance for lubrication and cooling of the engine may increase the efficiency of the system due to the reduction in mass, as well as the advantage of cooling the engine with a substance rather than relying on air cooling, which can be problematic throughout the engine.

[0077] Some embodiments of electric engines may include various components for monitoring flammable fluids and preventing flammable materials from entering specific sections of the electric engine. Some embodiments may include an electric engine having a wet zone enclosure which may be defined by a gearbox, motor, and / or heat exchanger. In some embodiments, an electric engine may have up to 4 liters or more of air in the motor gearbox housing in contact with the engine oil. Embodiments of the motor gearbox housing may use a breather to equalize internal and external pressures. Embodiments of the breather may include protruding over nearby design features to prevent accidental ingress of external fluids. Some embodiments may include a breather having a screen and a bypass entry path to prevent ingress of external debris. Embodiments may include a sight glass present in both tilt and rise electric engines to ensure that the oil is not overfilled or underfilled during maintenance.

[0078] Some embodiments of the electric engine may include, as necessary, active protection features in the forward and rear electric engines, such as monitoring vibrations throughout the engine, as well as internal temperatures throughout the engine, including oil temperature, stator winding set temperature, inverter bulk capacitor temperature, power module temperature, control board power module temperature, control board control processor temperature, control board monitor processor temperature, internal hotspot temperature, and various other operating conditions. Such monitoring may be achieved using various sensors positioned 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 the self-ignition temperature of the fluid. In some embodiments, various sensors used to monitor the operating conditions throughout the engine may report the operating conditions to the flight control system. Some embodiments may include threshold operating values ​​that may be required before the operating values ​​are 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 the detection of operating conditions. Some embodiments may include reducing the amount of power supplied to the electric propulsion system and / or reducing power in order to reduce mechanical wear or friction sparks from vibration, in order to reduce the temperature of components present in the electric propulsion system. Furthermore, some embodiments may include reducing power to an electric propulsion system where the detected efficiency of the inverter is below the target efficiency. In some embodiments, for example, if there are 12 electric propulsion systems 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 in order to counteract the reduction in lift generated by one electric propulsion system. In some embodiments, the flight control system may establish various thresholds for operating states in response to a decrease or increase in power to the electric propulsion systems.

[0079] Some embodiments may include a high-voltage power system that may have fuses at high-voltage battery terminals that can quickly and irreversibly disconnect engine electrical connections to mitigate and avoid overcurrent events. Such overcurrent protection may be activated when the current consumption of the electric engine is greater than the overcurrent operation. Thus, in some embodiments, fault conditions leading to overcurrent may only lead to transient overheating, arcing, or sparking faults. Some embodiments may include a fire threat characterization test ignition source that may be selected to be a more severe ignition source than a short circuit that occurs within the electric engine and is opened by the engine fuse. In some embodiments, the inverter may be able to detect AC overcurrents, isolate misphases, and / or continuously monitor the input DC voltage and apply protective actions to maintain the voltage below the overvoltage operating limit.

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

[0081] Some elements can generate high heat loads only during specific operating periods. For example, some lift propellers can only be used during takeoff, landing, and hovering, and can be shut off during cruising. Therefore, such lift propellers can generate high heat loads during takeoff, landing, and hovering, and generate little to no heat during cruising.

[0082] B. Exemplary Inverter Embodiments Figure 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 Figure 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 a 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 in response to a first pulse-width modulation (PWM) vector to drive the stator windings of a first set of electric motors M1. The second inverter circuit 1120 is configured to respond to the second PWM vector by converting 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 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 and second PWM vectors may be 0.25%, 0.5%, 1%, or 2% or less of the switching cycle period. For example, the delay may be within 50 nanoseconds. Therefore, the first inverter circuit 1110 is configured to output a first set of three-phase AC voltages (e.g., u1, v1, w1), and the second inverter circuit 1120 is configured to output a second set of three-phase AC voltages (e.g., u2, v2, w2), and the phase of the first set of three-phase AC voltages and the corresponding phase of the second set of three-phase AC voltages are two interleaved phases with a phase shift of substantially 180 degrees (e.g., plus or minus 5 degrees).

[0083] In particular, in the electric propulsion system 1100, two inverter circuits 1110 and 1120 are electrically coupled to an internal high-voltage DC supply bus and are configured to drive a dual three-phase motor M1 by providing corresponding three-phase AC voltages u1, v1, and w1 and three-phase AC voltages u2, v2, and w2. 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 configured to drive motor M1 by converting the bus voltage Vbus on a high-voltage DC supply bus into three-phase AC power. When inverter circuits 1110 and 1120 convert DC power to AC power, there is a voltage difference between the power supply and the neutral point of the load, which is called the common-mode voltage. Common-mode current resulting from the common-mode voltage in the inverter can be detrimental to the electrical system. Specifically, common-mode voltage can lead to motor failure, premature bearing failure, control device glitches, etc. In response to reducing common-mode noise, filter components may be installed in the electric propulsion system 1100.

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

[0086] For example, a DC common-mode filter 1130 may be located between a DC power supply 1160 and a 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 such 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, larger filters add volume and mass, which can lead to further losses.

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

[0088] The winding arrangement of the motor M1 applied in the electric propulsion system 1100 may differ in various embodiments. For example, the phase difference of the electrical angles between the two sets of three-phase windings may be designed to reduce harmonic components. In some embodiments, the stator windings of the first set of motor M1 and the stator windings of the second set are shifted by substantially 180 degrees (e.g., plus or minus 5 degrees). That is, the motor phase matching between independent winding sets may be out of phase by about 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 operating modes, the common-mode voltage can be attenuated via equal and opposite PWM vectors.

[0090] Figure 12A illustrates 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 illustrates 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, consistent with some embodiments of the present disclosure, in one cycle.

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

[0092] During operation, the switches in inverter circuits 1110 and 1120 are controlled so that both switches in the same leg (i.e., the upper and lower switches) do not turn on simultaneously to avoid a short circuit in the DC power supply. This can be achieved by the 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. Thus, the switching vectors SV0 to SV7 include six active switching vectors SV1 to SV6 and two zero vectors SV0 and SV7.

[0093] As shown in Figure 12A, the switching vector SV0 = {000} represents the upper switches of the three phases U, V, and W being off, while the lower switches of the three phases U, V, and W are on. The switching vector SV1 = {100} represents the upper switch of the U phase being on, and the upper switches of the V and W phases being off. The switching vector SV2 = {110} represents the upper switches of the U and V phases being on, and the upper switch of the W phase being off. The switching vector SV3 = {010} represents the upper switch of the V phase being on, and the upper switches of the U and W phases being off. The switching vector SV4 = {011} represents the upper switches of the V and W phases being on, and the upper switch of the U phase being off. The switching vector SV5 = {001} represents the upper switch of the W phase being on, and the upper switches of the U and V phases being off. Switching vector SV6 = {101} indicates that the upper switches for the U and W phases are on, and the upper switch for the V phase is off. Switching vector SV7 = {111} indicates that the upper switches for the three phases U, V, and W are on, while the lower switches for 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 reference SVM starting from switching vector SV0 with the upper switches of the three phases U, V, and W off, and the second inverter circuit 1120 is controlled according to an inverted midpoint reference SVM starting from switching vector SV7 with the upper switches of the three phases U, V, and W on. Thus, as shown in Figure 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, for the first inverter circuit 1110, the resulting PWM vector V INV 1 and the PWM vector V for the second inverter circuit 1120 INV2 and are equal and opposite PWM vectors. Therefore, the common-mode voltage V resulting from the first inverter circuit 1110 is 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 reduction of the common-mode voltage, the total common-mode voltage V SYS-CM It is zero.

[0095] In some embodiments, 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 suppression between any number of phases may be implemented so that the combination of PWM vectors can suppress 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 a power source (e.g., from a DC bus capacitor 1170), and the first and second inverter circuits 1110 and 1120 operate according to PWM signals that have equal duty cycles but are substantially out of phase by 180 degrees.

[0096] Common-mode voltage suppression achieved by the electric propulsion system 1100 having two inverters can reduce the volume and mass of the electromagnetic compatibility (EMC) components required in the system. In some embodiments, common-mode voltage suppression can provide a noise level at least 30-40 dB lower 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] Refer to Figure 13A illustrating an inverter circuit 1300 for motor control in an electric propulsion system, consistent with some embodiments of the present disclosure. The inverter circuit 1300, which is a voltage source inverter (VSI), may use FETs S1-S6 as switches. FETs S1-S6 can be selectively turned on or off in response to control signals from a controller following the SVM algorithm described above to generate a three-phase AC output voltage for driving the motor M1.

[0098] As shown in the embodiment of Figure 13A, the inverter circuit 1300 includes a bus capacitor 1320 configured to stabilize a DC bus voltage Vbus, and switches (e.g., FETs S1 to S6) forming a plurality of phase legs. Each phase leg includes an upper switch (e.g., FETs S1, S3, or S5) positioned between the positive terminal of the bus capacitor 1320 and the AC output terminal of the phase leg, and a lower switch (e.g., FETs S2, S4, or S6) positioned between the negative terminal of the bus capacitor 1320 and the AC output terminal of the phase leg.

[0099] In the inverter circuit 1300, FETs S1 to S6 can operate independently. If any of FETs S1 to S6 malfunction during operation, a single-phase short-circuit fault may occur if one leg is short-circuited, potentially generating an uncontrolled current through the corresponding phase as the machine rotates. This type of fault can lead to a fire hazard, create resistance to the system, impose very high torque pulsations on the system, or any other potential hazards. Therefore, it would be beneficial to reduce the torque present in response to a single-phase short-circuit fault during system operation.

[0100] The inverter circuit 1300 includes a first discharge circuit 1310 between the 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 the energy stored in the bus capacitor 1320 (i.e., the bulk capacitor) when a switch 1314 is closed in response to a corresponding command signal from a 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-circuit the bus capacitor 1320 in response to the DC bus voltage Vbus being below a fault condition threshold. Alternatively, during the discharge of the bus capacitor 1320, when the bus voltage Vbus is below a desired safety threshold, a bridge short-through of all FETs S1-S6 can be applied to short-circuit the HV bus, ensuring a safe discharge process and distributing the heat generated during a single-phase fault across all FETs S1-S6 in the inverter circuit 1300, thereby ensuring safety and reducing damage in the event of a fault.

[0102] In addition, as shown in Figure 13A, pyrofuses F1 and F2 are coupled between the inverter circuit 1300 and the DC voltage source Vin (e.g., a battery pack). For example, pyrofuses F1 and F2 may be types of fuses configured to be actuated by an external source when circuit disconnection and isolation are required. For example, the 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. The 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] Refer to Figure 13B illustrating another inverter circuit 1300 for motor control in an electric propulsion system, consistent with some embodiments of the present disclosure. The inverter circuit 1300 in Figure 13B is also a voltage source inverter (VSI) that uses FETs S1-S6 as switches to output three-phase AC output voltages to drive a motor M1.

[0104] In comparison with the inverter circuit 1300 in Figure 13A, the inverter circuit 1300 in Figure 13B further includes a comparator circuit 1330 and a second discharge circuit 1340 that form another discharge path across the bus capacitor 1320.

[0105] In particular, in the embodiment shown in Figure 13B, the comparator circuit 1330 is configured to monitor the bus voltage Vbus 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 a 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 verify that a power source (e.g., a battery) has been disconnected from the HV DC bus.

[0106] For example, 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 Vbus gradually decreases, and a voltage V1, which is a constant fraction of the bus voltage Vbus, also decreases. When voltage V1 is lower than the reference voltage Vref, it is determined that the bus voltage Vbus is lower than a predetermined threshold, indicating that the power supply 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 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 both the command signal Cmd and the signal output by the comparator 1334 are 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 rapid discharge component 1342 may be, but is not limited to, a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance device, or any combination thereof. The rapid discharge component 1342 may be any other component capable of achieving rapid discharge to remove residual energy on the HV DC bus.

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

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

[0111] Step 1410 detects whether a fault has occurred in one of the switches (e.g., FETs S1 to S6) in the inverter circuit 1300. In some embodiments, the inverter circuit 1300 is configured to determine whether a single-phase short-circuit fault has occurred in any one of the FETs S1 to S6. For example, the 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 or current signal within the inverter circuit 1300.

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

[0113] After a three-phase short circuit is applied, in step 1430, the inverter circuit 1300 is disconnected from the power supply in response to the detection of a single-phase short-circuit fault. In some embodiments, the inverter circuit 1300 is configured to instruct the battery management system (BMS) in the system so that a protection mechanism can 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, in order to protect against fault conditions for safe operation.

[0114] In some embodiments, step 1430 includes steps 1432, 1434, and 1436. For example, in step 1432, the inverter circuit 1300 may send a command 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 the inverter circuit 1300 from the power supply.

[0115] As an alternative method, 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. In addition, 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 was activated.

[0116] In other words, 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 command signals to the corresponding pyrofuse drivers to activate the pyrofuses, and thus electrically isolate the battery pack from the connected inverter circuit 1300 when other types of faults occur that should be protected from overcurrent. The inverter circuit 1300 may then, in response to receiving an acknowledgment signal from the BMS, 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 supply. In some embodiments, after receiving an acknowledgment 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 switch 1314 and use discharge resistor 1312 to form a discharge path for discharging energy. Thus, the bus voltage Vbus across the bus capacitor 1320 gradually decreases during the discharge process. In some embodiments, in step 1440, the inverter circuit 1300 may perform rapid discharge by using multiple discharge circuits in parallel.

[0118] For example, in step 1442, in response to confirmation that the inverter circuit 1300 has been disconnected from the power supply, the first switch (e.g., switch 1314 in Figure 13B) in the first discharge circuit (e.g., first discharge circuit 1310 in Figure 13B) is closed to provide the first discharge path. In step 1444, the bus voltage Vbus across the bus capacitor 1320 is monitored by the comparator circuit (e.g., comparator circuit 1330 in Figure 13B). In step 1446, in response that the bus voltage Vbus is below a second threshold, the second switch (e.g., switch 1344 in Figure 13B) in the second discharge circuit (e.g., second discharge circuit 1340 in Figure 13B) is closed to provide the second discharge path in parallel with the first discharge path.

[0119] In step 1450, in response to the bus voltage Vbus being below a first threshold, several switches in the inverter circuit 1300 (e.g., FETs S1-S6 in Figure 13B) are closed to short-circuit the bus capacitor 1320. In some embodiments, the inverter circuit 1300 is configured to detect the bus voltage Vbus across the bus capacitor 1320, and when the bus voltage Vbus is below a certain threshold voltage (e.g., about 50V), the inverter circuit 1300 may apply a bridge short-through to short-circuit all six FETs S1-S6, thereby short-circuiting the high-voltage bus. The threshold voltage can be designed based on practical needs. For example, the threshold voltage range may be about 40V-60V. Since the short-circuit configuration is activated after the bus voltage Vbus drops below a certain level during the rapid discharge process, the resulting current will be within a safe current limit and will not damage the components in the inverter circuit 1300 (e.g., FETs S1-S6). After the high-voltage bus is short-circuited, the motor M1 is stopped accordingly.

[0120] The operation of Method 1400 described above allows heat generated during a single-phase fault to be distributed among all FETs in the inverter circuit 1300, ensuring safety and reducing damage when a fault occurs. Furthermore, any one or more steps 1410-1450 performed in Method 1400 may occur within 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 the inverter circuit 1300. The control circuit may also provide pyro signals to activate one or both of the pyrofuses F1 and F2. In some other embodiments, the system may also include multiple control circuits for transmitting pyro signals to activate pyrofuses F1 and F2 and control signals to control one or more of the FETs in the 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 specific stages. The step(s) may be performed throughout Method 1400 as needed. In addition, the above examples include exemplary steps and / or operations, which may be added, replaced, modified in order, and / or excluded as appropriate, without departing from the spirit and scope of this disclosure. The methods for controlling inverter circuits and the inverter circuits disclosed in various embodiments may also be used in a variety of fields or systems, including, but not limited to, automobiles, hybrid and electric vehicles, and electric motors.

[0121] Various embodiments of this specification are described in one aspect in the general context of steps or processes of a method that may be implemented by an integrated circuit including a circuit mechanism for carrying out a method for controlling an inverter circuit. The circuit mechanism may be configured to carry out the steps or processes of the method described above. For example, the circuit mechanism may include one or more controllers, one or more processors, or a combination thereof, for controlling the inverter circuit disclosed in various embodiments of this disclosure.

[0122] Various embodiments of this specification are described in the general context of method steps or processes that may be implemented by computer program products embodied in temporary or non-temporary computer-readable media storing computer-executable instructions, such as program code, which are executed by one or more processors or one or more controllers in a system. Computer-readable media may include, but are not limited to, read-only memory devices (ROM), random-access memory devices (RAM), compact discs (CDs), digital versatile discs (DVDs), and removable and non-removable storage devices.

[0123] Generally, a program module may include routines, programs, objects, components, data structures, etc., that perform a specific task or implement a specific abstract data type. Computer executable instructions, associated data structures, and program modules represent examples of program code for performing steps of the methods disclosed herein. A particular sequence of such executable instructions or associated data structures represents an example of corresponding behavior for implementing the functionality described in such steps or processes.

[0124] Refer to Figure 15, a graph showing the bus voltage against time for an HV DC bus during a discharge period, consistent with some embodiments of the present disclosure. As shown in curve 1510, in a conventional discharge process, the voltage, current, and charge can decay exponentially during capacitor discharge. On the other hand, as shown in curve 1520, some embodiments of the present disclosure can enable rapid discharge. Thus, initially, the bus voltage Vbus can decay exponentially along an exponential decay curve, but the bus voltage Vbus may decay at a certain threshold V at time T. TWhen it drops, the bus voltage Vbus may drop rapidly to zero instead of continuing along an exponential decay curve. As described in the above embodiments, when the inverter circuit 1300 detects at time T that the bus voltage Vbus is lower than a specific threshold V T when it is lower than T , various circuit mechanisms (e.g., the rapid discharge component 1342 in FIG. 13B) can be used to rapidly discharge the remaining voltage.

[0125] Refer to FIGS. 16A - 16D, which are graphs showing EMI noise consistent with some embodiments of the present disclosure. FIGS. 16A - 16B are graphs showing the EMI noise detected on the positive and negative sides of a dual - inverter system, respectively, consistent with some embodiments of the present disclosure. FIGS. 16C - 16D are graphs showing the EMI noise detected on the positive and negative sides of a single - inverter system, respectively, consistent with some embodiments of the present disclosure. As shown in FIGS. 16A - 16D, in various embodiments disclosed in the present disclosure, the EMI noise in a 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, by the common - mode noise suppression / killing achieved by an electric propulsion system having two inverters, the volume and mass of the required electromagnetic compatibility (EMC) components can be reduced. As a result, a lightweight engine design can be realized.

[0126] In the above specification, the embodiments have been described with reference to numerous specific details that may vary depending on the implementation aspect. Specific adaptations and modifications of the described embodiments can be made. Also, the sequence of steps shown in the figures is for illustrative purposes only and is not intended to be limited to any particular sequence of steps. Therefore, those skilled in the art can understand that these steps can be implemented in a different order while implementing the same method.

[0127] As used herein, unless otherwise stated, the term “or” encompasses all possible combinations, except in cases where it is impossible to achieve. For example, if it is stated that a module may include A or B, then unless otherwise stated or impossible to achieve, the module may include A, or B, or A and B. As a second example, if it is stated that a module may include A, B, or C, then unless otherwise stated or impossible to achieve, the module may include A, or B, or C, or A and B, or A and C, or B and C, or A and B and C.

[0128] Exemplary embodiments are disclosed in the drawings and specification. It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed apparatus, systems, and related methods. Other embodiments will be apparent to those skilled in the art from the considerations herein and from the practice of the disclosed apparatus, systems, and related methods. The specification and examples are for illustrative purposes only, and the true scope is intended to be shown by the following claims and their equivalents.

[0129] Embodiments may be further described using the following clauses. Clause Set 1 1. A propulsion system for aircraft, An electric motor configured to drive one or more propellers of an aircraft, A capacitor configured to stabilize the DC bus voltage, A first inverter circuit, coupled to a capacitor and configured to convert a DC bus voltage on a first bus of the first inverter circuit to an AC voltage based on a first pulse-width modulation (PWM) vector, to drive the stator windings of a first set of electric motors, A propulsion system comprising: a second inverter circuit coupled to a 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, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors. 2. The propulsion system described in Clause 1, wherein the stator windings of the first set and the stator windings of the second set are shifted substantially by 180 degrees. 3. The propulsion system according to Clause 1 or 2, wherein the first inverter circuit is controlled using midpoint reference space vector modulation. 4. A propulsion system according to any one of the preceding clauses, wherein the second inverter circuit is controlled using inverted midpoint reference space vector modulation. 5. The propulsion system according to any one of the preceding clauses, wherein a first inverter circuit is configured to output a first set of three-phase AC voltages, and a second inverter circuit is configured to output a second set of three-phase AC voltages. 6. The propulsion system according to Clause 5, wherein the phases of the three-phase AC voltages of the first set and the corresponding phases of the three-phase AC voltages of the second set are two interleaved phases having a substantially 180-degree phase shift. 7. The propulsion system according to any one of the preceding clauses, further comprising a DC common-mode filter coupled to a capacitor and configured to reduce common-mode signals on the DC side of the first inverter circuit and the second inverter circuit. 8. The propulsion system according to 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 for reducing common-mode signals. Clause Set 2 9. A method for controlling a propulsion system for an aircraft, The capacitor stabilizes the DC bus voltage, A first inverter circuit coupled to a capacitor converts the DC bus voltage to an AC voltage according to a first pulse-width modulation (PWM) vector to drive the stator windings of a first set of electric motors, A second inverter circuit coupled to a capacitor converts the DC bus voltage to an AC voltage in response to a second PWM vector, thereby driving the stator windings of a second set of electric motors, wherein the first PWM vector and the second PWM vector are substantially equal and opposite vectors. A method comprising driving one or more propellers of an aircraft with an electric motor. 10. The method according to Clause 9, wherein the stator windings of the first set and the stator windings of the second set are shifted by substantially 180 degrees. 11. The method according to clause 9 or 10, further comprising controlling the first inverter circuit using midpoint reference space vector modulation. 12. The method according to any one of the clauses 9 to 11, further comprising controlling a second inverter circuit using inverted midpoint reference space vector modulation. 13. The first inverter circuit outputs a three-phase AC voltage for the first set to drive the stator windings of the first set, The method according to any one of the clauses 9 to 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 according to clause 13, wherein the phase of the three-phase AC voltage of the first set and the corresponding phase of the three-phase AC voltage of the second set are two interleaved phases having a substantially 180-degree phase shift. 15. The method according to any one of the claims 9 to 14, further comprising reducing the common-mode signal on the DC side of the first inverter circuit and the second inverter circuit by a DC common-mode filter coupled to a capacitor. 16. The method according to any one of the claims 9 to 15, further comprising reducing the common-mode signals on the AC sides of the first inverter circuit and the second inverter circuit by one or more AC common-mode chokes coupled to the AC side of the first inverter circuit or the second inverter circuit. Clause Set 3 17. An integrated circuit comprising a circuit mechanism for performing a method for controlling a propulsion system for an aircraft, wherein the circuit mechanism is A capacitor-coupled first inverter circuit is controlled to convert the DC bus voltage to an AC voltage according to a first pulse-width modulation (PWM) vector, thereby driving the stator windings of a first set of electric motors. An integrated circuit configured to control a capacitor-coupled second inverter circuit to drive a second set of stator windings of an electric motor by converting a DC bus voltage to an AC voltage in response to a second PWM vector, wherein the first and second PWM vectors are substantially equal and opposite vectors for driving one or more propellers of an aircraft by an electric motor. 18. The integrated circuit described in Clause 17, wherein the stator windings of the first set and the stator windings of the second set are shifted substantially by 180 degrees. 19. The integrated circuit according to Clause 17 or 18, wherein the circuit mechanism is further configured to control the first inverter circuit using midpoint reference space vector modulation. 20. An integrated circuit according to any one of the clauses 17 to 19, wherein the circuit mechanism is further configured to control a second inverter circuit using inverted midpoint reference space vector modulation. 21. The circuit mechanism is further, The first inverter circuit is controlled to output a first set of three-phase AC voltages to drive the first set of stator windings. An integrated circuit according to any one of clauses 17 to 20, configured to control a second inverter circuit so as to output a second set of three-phase AC voltages to drive a second set of stator windings. 22. The integrated circuit according to Clause 21, wherein the phases of the three-phase AC voltages of the first set and the corresponding phases of the three-phase AC voltages of the second set are two interleaved phases having a substantially 180-degree phase shift. 23. An integrated circuit according to any one of the clauses 17 to 22, wherein a DC common-mode filter is coupled to a 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 a first inverter circuit or a second inverter circuit in order to reduce common-mode signals. Clause Set 4 25. An inverter circuit for an aircraft propulsion system, A capacitor configured to stabilize the DC bus voltage, A plurality of switches forming a plurality of phase legs, wherein at least one of the phase legs includes an upper switch positioned between the positive terminal of a capacitor and the AC output terminal of the phase leg, and a lower switch positioned between the negative terminal of a capacitor and the AC output terminal of the phase leg, The system comprises a first discharge circuit connected in parallel to the capacitor and configured to provide a first discharge path for discharging the energy stored in the capacitor, An inverter circuit in which multiple switches are controlled to short-circuit a capacitor in response to the DC bus voltage being lower than a first threshold in a fault condition associated with the inverter circuit. 26. The inverter circuit according to Clause 25, wherein the first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and the inverter circuit is configured to discharge a capacitor by closing the first switch in response to a single-phase short circuit. 27. The inverter circuit according to clause 25 or 26, further comprising a second discharge circuit coupled in parallel with the capacitor and configured to provide a second discharge path for discharging energy stored in the capacitor. 28. The inverter circuit according to Clause 27, wherein the second discharge circuit comprises a discharge element connected in series and a second switch. 29. An inverter circuit as described in Clause 28, wherein the discharge element includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance device, or any combination thereof. 30. An inverter circuit according to any one of clauses 25 to 29, further comprising a comparator circuit configured to monitor the DC bus voltage across a capacitor and determine whether to discharge the energy stored in the capacitor through a second discharge circuit. 31. The comparator circuit is A resistor voltage divider comprising resistors connected in series to provide a first voltage, wherein the first voltage is a constant fraction of the DC bus voltage, A comparator, which is coupled to a resistor voltage divider, and The first voltage output from the resistor voltage divider is compared with the reference voltage. An inverter circuit according to clause 30, comprising a comparator configured to output an output signal in response to a first voltage being lower than a reference voltage. 32. The comparator circuit, A logic circuit, which is coupled to a comparator, and The output signal from the comparator and the command signal from the control circuit are received. The inverter circuit according to Clause 31, further comprising a logic circuit configured to output a control signal for selectively turning on a second switch of a second discharge circuit according to an output signal and a command signal. Clause Set 5 33. A method for controlling an inverter circuit, To detect whether a failure has occurred in one of the multiple switches in the inverter circuit, In response to the detection of a single-phase short-circuit fault, the inverter circuit is disconnected from the power supply, The first discharge circuit provides a first discharge path for discharging the DC bus voltage across the capacitor of the inverter circuit after the inverter circuit has been disconnected from the power supply. A method comprising controlling multiple switches in an inverter circuit to short-circuit a capacitor in response to a DC bus voltage being lower than a first threshold. 34. The method according to 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 comprises discharging a capacitor by closing the first switch in response to a single-phase short circuit via an inverter circuit. 35. The method according to 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 with the capacitor. 36. The inverter circuit according to Clause 35, wherein the second discharge circuit comprises a discharge element connected in series and a second switch. 37. An inverter circuit according to Clause 36, wherein the discharge element includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance device, or any combination thereof. 38. The method according to any one of the clauses 33 to 37, further comprising monitoring the DC bus voltage across the capacitor using 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, wherein the first voltage is a constant fraction of the bus voltage. The method according to any one of the clauses 33 to 38, further comprising comparing a first voltage with a reference voltage using a comparator, and outputting an output signal in response that the first voltage is lower than the reference voltage. 40. The logic circuit receives the output signal from the comparator and the command signal from the control circuit. The method according to Clause 39, further comprising: a logic circuit outputting a control signal for selectively turning on a second switch in a second discharge circuit according to an output signal and a command signal. Clause Set 6 41. An integrated circuit comprising a circuit mechanism for carrying out a method for controlling an inverter circuit, wherein the circuit mechanism is a method, To detect whether a failure has occurred in one of the multiple switches in the inverter circuit, In response to the detection of a single-phase short-circuit fault, the inverter circuit is disconnected from the power supply, The inverter circuit is controlled by a first discharge circuit to provide a first discharge path for discharging the DC bus voltage across the capacitors of the inverter circuit after the inverter circuit has been disconnected from the power supply. An integrated circuit configured to perform a method including controlling an inverter circuit to short-circuit a capacitor by controlling a plurality of switches in the inverter circuit in response to a DC bus voltage being lower than a first threshold. 42. The integrated circuit according to Clause 41, wherein the first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and the circuit mechanism is configured to control an inverter circuit to discharge a capacitor by closing the first switch in response to a single-phase short circuit. 43. The circuit mechanism is The integrated circuit according to Clause 41 or 42, configured to control an inverter circuit to provide a second discharge path for discharging a DC bus voltage by a second discharge circuit coupled in parallel with the capacitor. 44. The integrated circuit according to Clause 43, wherein the second discharge circuit comprises a discharge element connected in series and a second switch. 45. An integrated circuit as described in Clause 44, wherein the discharge element includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance device, or any combination thereof. 46. ​​An integrated circuit according to any one of clauses 43 to 45, wherein a comparator circuit monitors the DC bus voltage across the capacitor and determines whether to discharge the energy stored in the capacitor through a second discharge circuit. 47. The DC bus voltage across the capacitor is A resistor divider comprising resistors connected in series provides a first voltage, wherein the first voltage is a constant fraction of the bus voltage. An integrated circuit according to any one of clauses 41 to 46, which 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. The DC bus voltage across the capacitor is The logic circuit receives the output signal from the comparator and the command signal from the control circuit. An integrated circuit as described in Clause 47, which is monitored by a logic circuit to output a control signal for selectively turning on a second switch in a second discharge circuit according to an output signal and a command signal. Clause Set 7 49. An inverter circuit for an aircraft propulsion system, A capacitor configured to stabilize the DC bus voltage, A plurality of switches forming a plurality of phase legs, wherein at least one of the phase legs includes an upper switch positioned between the positive terminal of a capacitor and the AC output terminal of the phase leg, and a lower switch positioned between the negative terminal of a capacitor and the AC output terminal of the phase leg, A first discharge circuit is connected in parallel to the capacitor and configured to provide a first discharge path for discharging the energy stored in the capacitor, An inverter circuit comprising: a second discharge circuit coupled in parallel with a 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 according to Clause 49, wherein the first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and the inverter circuit is configured to discharge a capacitor by closing the first switch in response to a single-phase short circuit. 51. The inverter circuit according to Clause 49 or 50, wherein the second discharge circuit comprises a discharge element connected in series and a second switch. 52. An inverter circuit as described in Clause 51, wherein the discharge element includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance device, or any combination thereof. 53. An inverter circuit according to any one of clauses 49 to 52, further comprising a comparator circuit configured to monitor the DC bus voltage across a capacitor and determine whether to discharge the energy stored in the capacitor through a second discharge circuit. 54. The comparator circuit is A resistor voltage divider comprising resistors connected in series to provide a first voltage, wherein the first voltage is a constant fraction of the DC bus voltage, A comparator, which is coupled to a resistor voltage divider, and The first voltage output from the resistor voltage divider is compared with the reference voltage. An inverter circuit according to Clause 53, comprising a comparator configured to output an output signal in response to a first voltage being lower than a reference voltage. 55. The comparator circuit, A logic circuit, which is coupled to a comparator, and The output signal from the comparator and the command signal from the control circuit are received. The inverter circuit according to Clause 54, further comprising a logic circuit configured to output a control signal for selectively turning on a second switch of a second discharge circuit according to an output signal and a command signal. Clause Set 8 56. A method for controlling an inverter circuit, To detect whether a failure has occurred in one of the multiple switches in the inverter circuit, In response to the detection of a single-phase short-circuit fault, the inverter circuit is disconnected from the power supply, After the inverter circuit is disconnected from the power supply, the bus voltage across the bus capacitor of the inverter circuit is discharged, In response to confirmation that the inverter circuit has been disconnected from the power supply, a first discharge circuit is used to provide a first discharge path. A method comprising providing a second discharge path in parallel with a first discharge path using a second discharge circuit in response to the bus voltage being lower than a threshold. 57. The first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and the bus voltage across the capacitor is discharged. The method according to Clause 56, further comprising closing a first switch in a first discharge circuit to provide a first discharge path in response to confirmation that the inverter circuit has been disconnected from the power supply. 58. The second discharge circuit comprises a discharge element and a second switch connected in series, and discharges the bus voltage across the capacitor. The method according to clause 56 or 57, further comprising closing a second switch in a second discharge circuit in response to a bus voltage being below a threshold to provide a second discharge path in parallel with the first discharge path. 59. The method according to Clause 58, wherein the discharge element includes 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 the capacitor is The method according to any one of the clauses 56 to 59, further comprising monitoring the bus voltage across the capacitor using 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 the capacitor is A resistor divider comprising resistors connected in series provides a first voltage, wherein the first voltage is a constant fraction of the bus voltage. The method according to any one of the claims 56 to 60, further comprising comparing a first voltage with a reference voltage using a comparator, and outputting an output signal in response that the first voltage is lower than the reference voltage. 62. Discharging the bus voltage across the capacitor is The logic circuit receives the output signal from the comparator and the command signal from the control circuit. The method according to Clause 61, further comprising: a logic circuit outputting a control signal for selectively turning on a second switch in a second discharge circuit according to an output signal and a command signal. Clause Set 9 63. An integrated circuit comprising a circuit mechanism for carrying out a method for controlling an inverter circuit, wherein the circuit mechanism is It detects whether a failure has occurred in one of the multiple switches in the inverter circuit. In response to the detection of a single-phase short-circuit fault, the inverter circuit is disconnected from the power supply. In response to confirmation that the inverter circuit has been disconnected from the power supply, a first discharge circuit is used to provide a first discharge path. An integrated circuit configured to control an inverter circuit so that, after the inverter circuit is disconnected from the power supply, the bus voltage across the capacitor of the inverter circuit is discharged by using a second discharge circuit to provide a second discharge path in parallel with the first discharge path in response to the bus voltage being below a threshold. 64. The first discharge circuit comprises a discharge resistor and a first switch connected in series with the discharge resistor, and the inverter circuit is The integrated circuit according to Clause 63, which is controlled to discharge the bus voltage across a capacitor by closing a first switch in a first discharge circuit to provide a first discharge path in response to confirmation that the inverter circuit has been disconnected from the power supply. 65. The second discharge circuit comprises a discharge element and a second switch connected in series, and the inverter circuit is The integrated circuit according to clause 63 or 64, which controls the bus voltage across a capacitor to discharge by closing a second switch in a second discharge circuit to provide a second discharge path in response to the bus voltage being below a threshold. 66. The method according to Clause 65, wherein the discharge element includes a transient voltage suppression diode (TVS), a metal oxide varistor (MOV), a low-resistance device, or any combination thereof. 67. The bus voltage across the capacitor is The method according to any one of the clauses 63 to 66, wherein the bus voltage across the capacitor is monitored by a comparator circuit to determine whether to discharge the energy stored in the capacitor through a second discharge circuit. 68. The bus voltage across the capacitor is A resistor divider comprising resistors connected in series provides a first voltage, wherein the first voltage is a constant fraction of the bus voltage. An integrated circuit according to 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. The bus voltage across the capacitor is The logic circuit receives the output signal from the comparator and the command signal from the control circuit. An integrated circuit according to any one of clauses 63 to 68, which is discharged by outputting a control signal for selectively turning on a second switch in a second discharge circuit according to an output signal and a command signal via a logic circuit.

[0130] The embodiments disclosed herein are intended to be non-limiting. Those skilled in the art will understand that certain components and configurations of components can be modified without departing from the scope of the disclosed embodiments.

Claims

1. A propulsion system for aircraft, An electric motor configured to drive one or more propellers of the aircraft, A capacitor configured to stabilize the DC bus voltage, A first inverter circuit, coupled to the capacitor and configured to convert the DC bus voltage on the first bus of the first inverter circuit into an AC voltage based on a first pulse-width modulation (PWM) vector to drive the stator windings of a first set of electric motors, A second inverter circuit, coupled to the capacitor and configured to convert the DC bus voltage on the second bus of the second inverter circuit to an AC voltage based on a second PWM vector, thereby driving the stator windings of a second set of electric motors, is provided. The first PWM vector and the second PWM vector are substantially equal in magnitude and opposite in direction. The phase difference between the stator windings of the first set and the stator windings of the second set is substantially 180 degrees. Propulsion system.

2. The propulsion system according to claim 1, wherein the first inverter circuit is controlled using midpoint reference space vector modulation.

3. The propulsion system according to claim 1, wherein the second inverter circuit is controlled using inverted midpoint reference space vector modulation.

4. The propulsion system according to claim 1, wherein the first inverter circuit is controlled using midpoint-referenced space vector modulation, and the second inverter circuit is controlled using inverted midpoint-referenced space vector modulation.

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

7. The propulsion system according to claim 1, further comprising a DC common-mode filter 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.

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

9. 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. The phase of the three-phase AC voltage of the first set and the corresponding phase of the three-phase AC voltage of the second set are two interleaved phases having a substantially 180-degree phase shift. The propulsion system according to claim 1, further comprising a DC common-mode filter 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.

10. The first inverter circuit is controlled using midpoint-referenced space vector modulation, and the second inverter circuit is controlled using inverted midpoint-referenced space vector modulation. The propulsion system according to claim 1, further comprising a DC common-mode filter 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.

11. A method for controlling an aircraft propulsion system, The capacitor stabilizes the DC bus voltage, The first inverter circuit coupled to the capacitor converts the DC bus voltage to an AC voltage according to a first pulse-width modulation (PWM) vector to drive the stator windings of a first set of electric motors. The second inverter circuit coupled to the capacitor converts the DC bus voltage to an AC voltage in response to a second PWM vector, thereby driving the stator windings of a second set of electric motors. The first PWM vector and the second PWM vector are substantially equal in magnitude and opposite in direction. The phase difference between the stator windings of the first set and the stator windings of the second set is substantially 180 degrees, and the drive is performed. A method comprising driving one or more propellers of the aircraft with the electric motor.

12. The method according to claim 11, further comprising controlling the first inverter circuit using midpoint reference space vector modulation.

13. The method according to claim 11, further comprising controlling the second inverter circuit using inverted midpoint reference space vector modulation.

14. Controlling the first inverter circuit using midpoint reference space vector modulation, The method according to claim 11, further comprising controlling the second inverter circuit using inverted midpoint reference space vector modulation.

15. The first inverter circuit outputs a three-phase AC voltage for the first set to drive the stator windings of the first set, The method according to claim 11, further comprising: outputting a three-phase AC voltage of a second set by the second inverter circuit to drive the stator windings of the second set.

16. The method according to claim 15, wherein the phase of the three-phase AC voltage of the first set and the corresponding phase of the three-phase AC voltage of the second set are two interleaved phases having a substantially 180-degree phase shift.

17. The method according to claim 11, further comprising reducing the common-mode signal on the DC side of the first inverter circuit and the second inverter circuit by a DC common-mode filter coupled to the capacitor.

18. The method according to any one of claims 11 to 17, further comprising reducing the common-mode signal on the AC side of the first inverter circuit and the second inverter circuit by one or more AC common-mode chokes coupled to the AC side of the first inverter circuit or the second inverter circuit.

19. The first inverter circuit outputs a three-phase AC voltage for the first set to drive the stator windings of the first set, The second inverter circuit outputs a three-phase AC voltage for the second set to drive the stator windings of the second set, The invention further includes reducing the common-mode signal on the DC side of the first inverter circuit and the second inverter circuit by using a DC common-mode filter coupled to the capacitor, The method according to claim 11, wherein the phase of the three-phase AC voltage of the first set and the corresponding phase of the three-phase AC voltage of the second set are two interleaved phases having a substantially 180-degree phase shift.

20. Controlling the first inverter circuit using midpoint reference space vector modulation, Controlling the second inverter circuit using inverted midpoint reference space vector modulation, The method according to claim 11, further comprising reducing the common-mode signal on the DC side of the first inverter circuit and the second inverter circuit by a DC common-mode filter coupled to the capacitor.

21. An integrated circuit comprising a circuit mechanism for performing a method for controlling a propulsion system for an aircraft, wherein the circuit mechanism is A capacitor-coupled first inverter circuit is controlled to convert a DC bus voltage to an alternating current (AC) voltage according to a first pulse-width modulation (PWM) vector, thereby driving the stator windings of a first set of electric motors. The system is configured to control a second inverter circuit coupled to the capacitor in response to a second PWM vector, converting the DC bus voltage to an AC voltage to drive the stator windings of a second set of electric motors. The first PWM vector and the second PWM vector are substantially equal in magnitude and opposite in direction in order to drive one or more propellers of the aircraft with the electric motor. An integrated circuit in which the phase difference between the stator windings of the first set and the stator windings of the second set is substantially 180 degrees.

22. The integrated circuit according to claim 21, wherein the circuit mechanism is further configured to control the first inverter circuit using midpoint reference space vector modulation.

23. The integrated circuit according to claim 21, wherein the circuit mechanism is further configured to control the second inverter circuit using inverted midpoint reference space vector modulation.

24. The circuit mechanism is further configured to control the first inverter circuit using midpoint reference space vector modulation and to control the second inverter circuit using inverted midpoint reference space vector modulation. The integrated circuit according to claim 21.

25. The aforementioned circuit mechanism further, The first inverter circuit is controlled to output a three-phase AC voltage to the first set and drive the stator windings of the first set. The integrated circuit according to claim 21, configured to control the second inverter circuit so that it outputs a second set of three-phase AC voltages to drive the second set of stator windings.

26. The integrated circuit according to claim 25, wherein the phases of the three-phase AC voltages of the first set and the corresponding phases of the three-phase AC voltages of the second set are two interleaved phases having a substantially 180-degree phase shift.

27. The integrated circuit according to claim 21, 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.

28. The integrated circuit according to any one of claims 21 to 27, wherein one or more AC common-mode chokes are coupled to the AC side of the first inverter circuit or the second inverter circuit in order to reduce common-mode signals.

29. The aforementioned circuit mechanism further, The first inverter circuit is controlled to output a three-phase AC voltage to the first set and drive the stator windings of the first set. The second inverter circuit is configured to output a three-phase AC voltage to the second set and control the second inverter circuit to drive the stator windings of the second set. The phases of the three-phase AC voltages of the first set and the corresponding phases of the three-phase AC voltages of the second set are two interleaved phases having a substantially 180-degree phase shift. The integrated circuit according to claim 21, 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.

30. The circuit mechanism is further configured to control the first inverter circuit using midpoint reference space vector modulation and to control the second inverter circuit using inverted midpoint reference space vector modulation. The integrated circuit according to claim 21, 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.

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