Patent Grant
11912424
This disclosure relates generally to the field of powered aerial vehicles. More particularly, and without limitation, the present disclosure relates to innovations in aircrafts driven by electric propulsion systems. Certain aspects of the present disclosure generally relate to improvements in electric engines that may be used in aircrafts driven by electric propulsion systems and in other types of vehicles. Other aspects of the present disclosure generally relate to improvements in gearboxes that provide particular advantages in aerial vehicles and may be used in other types of vehicles.
The present disclosure addresses systems, components, and techniques primarily for use in a non-conventional aircraft driven by an electric propulsion system. For example, the tilt-rotor aircraft of the present disclosure may be configured for frequent (e.g., over 50 flights per work day), short-duration flights (e.g., less than 100 miles per flight) over, into, and out of densely populated regions. The aircraft may be configured to carry 4-6 passengers or commuters who have an expectation of a comfortable experience with low noise and low vibration. Accordingly, it may be desired that components of the aircraft are configured and designed to withstand frequent use without wearing, generate less heat and vibration, and that the aircraft include mechanisms to effectively control and manage heat or vibration generated by the components. Further, it may be intended that several of these aircraft operate near each other over a crowded metropolitan area. Accordingly, it may be desired that their components are configured and designed to generate low levels of noise interior and exterior to the aircraft, and to have a variety of safety and backup mechanisms. For example, it may be desired for safety reasons that the aircraft be propelled by a distributed propulsion system, avoiding the risk of a single point of failure, and that they are capable of conventional takeoff and landing on a runway. Moreover, it may be desired that the aircraft can safely vertically takeoff and land from and into relatively small or restricted spaces compared to traditional airport runways (e.g., vertiports, parking lots, or driveways) while transporting several passengers or commuters with accompanying baggage. These use requirements may place design constraints on aircraft size, weight, operating efficiency (e.g., drag, energy use), which may impact the design and configuration of the aircraft components.
Disclosed embodiments provide new and improved configurations of aircraft components that are not observed in conventional aircraft, and/or identified design criteria for components that differ from those of conventional aircraft. Such alternate configurations and design criteria, in combination addressing drawbacks and challenges with conventional components, yielded the embodiments disclosed herein for various configurations and designs of components for an aircraft driven by an electric propulsion system.
In some embodiments, the aircraft driven by an electric propulsion system of the present disclosure may be designed to be capable of both vertical and conventional takeoff and landing, with a distributed electric propulsion system enabling vertical flight, horizontal and lateral flight, and transition. Thrust may be generated by supplying high voltage electrical power to a plurality of electric engines of the distributed electric propulsion system, which may include the necessary components to convert the high voltage electrical 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 an electric engine connected to an onboard electrical power source, which may include a device capable of storing energy such as a battery or capacitor, and may include one or more systems for harnessing or generating electricity such as a fuel powered generator or solar panel array. Some disclosed embodiments provide for weight reduction and space reduction of components in the aircraft to, increase aircraft efficiency and performance. Disclosed embodiments also improve upon safety in passenger transportation using new and improved safety protocols and system redundancy in the case of a failure, to minimize any single points of failure in the aircraft propulsion system. Some disclosed embodiments also provide new and improved approaches to satisfying and exceeding aviation and transportation laws and regulations. For example, the Federal Aviation Administration enforces federal laws and regulations requiring safety components such as fire protective barriers adjacent to engines that use more than a threshold amount of oil or other flammable materials. A fire protective barrier may include an engine component or aircraft component designed, constructed, or installed with the primary purpose of being constructed so that no hazardous quantity of air, fluid, or flame can pass around or through the fire protective barrier and/or to protect against corrosion. In some embodiments, a fire protective barrier may include a component separate from additional components as recited herein. Persons of ordinary skill in the art would understand which components within an aircraft, including within an electric propulsion system, would act with the primary function of being a fire protective barrier. In some embodiments, a fire protective barrier may include a firewall, a fireproof barrier, a fire resistant barrier, a flame resistant barrier, or any other barrier capable of ensuring no hazardous quantity of air, fluid, or flame can pass around or through the barrier and/or to protect against corrosion. For example, while a fuselage may be constructed so that no hazardous quantity of air, fluid, or flame can pass around or through the fire protective barrier, and/or protect against corrosion, the fuselage may not be considered a fire protective barrier since the primary purpose of a fuselage is not to be a fire protective barrier. In some embodiments, electric propulsion systems provide for efficient and effective lubrication and cooling using less than the threshold level of oil, yielding an aircraft that does not require engine fire protective barriers, saving on aircraft weight while maximizing performance and efficiency.
In some embodiments, the distributed electric propulsion system may include twelve electric engines, which may be mounted on booms forward and aft of the main wings of the aircraft. A subset of the electric engines, such as those mounted forward of the main wings, may be tiltable mid-flight between a horizontally oriented position (e.g., to generate forward thrust for cruising) and a vertically oriented position (e.g., to generate vertical lift for takeoff, landing, and hovering). The propellers of the forward electric engines may rotate in a clockwise or counterclockwise direction. Propellers may counter-rotate with respect to adjacent propellers. The aft electric engines may be fixed in a vertically oriented position (e.g., to generate vertical lift). The propellers may also rotate in a clockwise or counterclockwise direction. In some embodiments, the difference in rotation direction may be achieved using the direction of engine rotation. In other embodiments, the engines may all rotate in the same direction, and gearing may be used to achieve different propeller rotation directions
In some embodiments, an aircraft may possess quantities of electric engines in various combinations of forward and aft engine configurations. For example, an aircraft may possess 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 where the number of forward electric engines and aft electric engines are not equivalent.
In some embodiments, for a vertical takeoff and landing (VTOL) mission, the forward and aft electric engines may provide vertical thrust during takeoff and landing. During flight phases where the aircraft is moving forward, the forward electric engines may provide horizontal thrust, while the propellers of the aft electric engines may be stowed at a fixed position in order to minimize drag. The aft electric engines may be actively stowed with position monitoring. Transition from vertical flight to horizontal flight and vice-versa may be accomplished via the tilt propeller subsystem. The tilt propeller subsystem may redirect thrust between a primarily vertical direction during vertical flight mode to a horizontal or near-horizontal direction during a forward-flight cruising phase. A variable pitch mechanism may change the forward electric engine's propeller-hub assembly blade collective angles for operation during the hover-phase, transition phase, and cruise-phase.
In some embodiments, in a conventional takeoff and landing (CTOL) mission, the forward electric engines may provide horizontal thrust for wing-borne take-off, cruise, and landing, and the wings may provide vertical lift. In some embodiments, the aft electric engines may not be used for generating thrust during a CTOL mission and the aft propellers may be stowed in place. In other embodiments, the aft electrical engines may be used at reduced power to shorten the length of the CTOL takeoff or landing.
In some embodiments, an electric engine for a vertical takeoff-and-landing aircraft may comprise an electric motor assembly including a stator and a rotor. Some embodiments may include an electric engine comprising an inverter assembly and a gearbox assembly including a sun gear. In some embodiments, an electric engine may include a main shaft including a length of the main shaft that extends from a first end of the main shaft through the gearbox assembly and through the electric motor assembly to a second end of the main shaft. In some embodiments, an electric engine may include a hydrodynamic bearing located between the main shaft and sun gear, and a bearing including an inner race mechanically coupled to the main shaft and an outer race mechanically coupled to the rotor. Some embodiments may include a bearing including an outer race mechanically coupled to an inner surface of the rotor.
The disclosed embodiments provide systems, subsystems, and components for new VTOL aircraft having various combinations of an electric propulsion system and cooling systems that maximize performance while minimizing weight.
In some embodiments, an electric propulsion system as described herein may generate thrust by supplying High Voltage (HV) electric power to an electric engine, which in turn converts HV power into mechanical shaft power which is used to rotate a propeller. An aircraft as described herein may include multiple electric engines mounted forward and aft of the wing. The engines may be mounted directly to the wing, or mounted to one or more booms attached to the wing. The amount of thrust each electric engine generates may be governed by a torque command from a Flight Control System (FCS) over a digital communication interface to each electric engine. Embodiments may include forward electric engines that are capable of altering their orientation, or tilt. Some embodiments include forward engines that may be a clockwise (CW) type or counterclockwise (CCW) type. The forward electric propulsion subsystem may consist of a multi-blade adjustable pitch propeller, as well as a variable pitch subsystem.
In some embodiments, an aircraft may include aft electric engines, or lifters, that can be of a clockwise (CW) type or counterclockwise (CCW) type. Some embodiments may include aft electric engines that utilize a multi-blade fixed pitch propeller.
As described herein, the orientation and use of the electric propulsion system components may change throughout the operation of the aircraft. In some embodiments, during vertical takeoff and landing, the forward propulsion systems as well as aft propulsion systems may provide vertical thrust during takeoff and landing. During the flight phases where the aircraft is in forward flight-mode, the forward propulsion systems may provide horizontal thrust, while the aft propulsion system propellers may be stowed at a fixed position to minimize drag. The aft electric propulsion systems may be actively stowed with position monitoring. Some embodiments may include a transition from vertical flight to horizontal flight and vice-versa. In some embodiments, the transitions may be accomplished via the Tilt Propeller System (TPS). The TPS reorients the electric propulsion system between a primarily vertical direction during vertical flight mode to a mostly horizontal direction during forward-flight mode. Some embodiments may include a variable pitch mechanism that may change the forward propulsion system propeller blade collective angles for operation during the hover-phase, cruise-phase and transition phase. Some embodiments may include a Conventional Takeoff and Landing (CTOL) configurations such that the tilters provide horizontal thrust for wing-borne take-off, cruise and landing phases. In some embodiments, the aft electric engines are not used for generating thrust during a CTOL mission and the aft propellers are stowed in place to minimize drag.
In some embodiments, an electric engine as described herein may possess design features which mitigate and protect against uncontained fire, such as utilizing non-hazardous quantity of flammable fluid contained in both the tilt and lift engines. For example, in some embodiments, the electric engine may be configured to utilize less than one quart of oil or another flammable fluid. Some embodiments may include the electric engine containing a non-hazardous quantity of air such that any fire may not be capable of maintaining a duration capable of migrating to another portion of the aircraft. In some embodiments, the non-hazardous quantity of air may be in contact with flammable liquids throughout the electric engine. Some examples may include an electric engine possessing up to one liter, two liters, three liters, four liters, five liters, ten liters, or twenty liters of air within the electric engine housing. In some embodiments, the amount of air present within the electric engine housing may possess a fixed ratio to the amount of oil, or other liquid for cooling, present within the electric propulsion system. Such a ratio may be driven by a determination of the sufficient thermal mass needed to properly cool the electric propulsion system. Some embodiments may include a ratio of about 3:1 of air to oil present within the electric propulsion system. Some embodiments may include an electric engine housing where 75% of the open volume, that is, interior volume that is not occupied by components of the electric engine, is comprised of air while 25% of the open volume is comprised of oil or some other liquid for cooling and/or lubricating. Some embodiments may also be configured without a nominal ignition source within the electric engines, possess an engine over temperature operating limit that may be more than 50° C. less than a flammable fluid auto-ignition temperature, possess overtemperature detection and protection, overvoltage detection and protection, and/or possess overcurrent detection and protection. Further, some embodiments may include an electric propulsion system where the bulk temperature of the electric propulsion system is lower than the autoignition temperature and flashpoint of the oil, or other liquid, present within the electric propulsion system in all normal operating conditions. In some embodiments, non-normal conditions that raise the bulk electric propulsion system temperature may result in system responses that prevent the exceedance of the oil, or other liquid, flashpoint and autoignition temperature. In some embodiments, the ratio of air to oil, or other liquid, may be such that if a fire were to occur, including if an arc were to cause a fire, within the electric engine housing, the amount of air present within the electric engine housing would not allow the fire to propagate to other areas of the aircraft. In some embodiments, these and other design features may yield an electric engine that is deemed by one or more guidelines or regulations to not be a designated fire zone.
Reference will now be made in detail to exemplary embodiments, examples of which are illustrated in the accompanying drawings. The following description refers to the accompanying drawings in which the same numbers in different drawings represent the same or similar elements unless otherwise represented. The implementations set forth in the following description of exemplary embodiments do not represent all implementations consistent with the disclosure. Instead, they are merely examples of apparatuses and methods consistent with aspects related to the subject matter recited in the appended claims.
In some embodiments, lift propellers 112, 212 may be configured for providing lift only, with all horizontal propulsion being provided by the tilt propellers. Accordingly, lift propellers 112, 212 may be configured with fixed positions and may only generate thrust during take-off, landing and hover phases of flight. Meanwhile, tilt propellers 114, 214 may be tilted upward into a lift configuration in which thrust from propellers 114, 214 is directed downward to provide additional lift.
For forward flight, tilt propellers 114, 214 may tilt from their lift configurations to their cruise configurations. In other words, the orientation of tilt propellers 114, 214 may be varied from an orientation in which the tilt propeller thrust is directed downward (to provide lift during vertical take-off, landing and hover) to an orientation in which the tilt propeller thrust is directed rearward (to provide forward thrust to aircraft 100, 200). The tilt propellers assembly for a particular electric engine may tilt about an axis of rotation defined by a mounting point connecting the boom and the electric engine. When the aircraft 100, 200 is in full forward flight, lift may be provided entirely by wings 104, 204. Meanwhile, in the cruise configuration, lift propellers 112, 212 may be shut off. The blades 120, 220 of lift propellers 112, 212 may be held in low-drag positions for aircraft cruising. In some embodiments, lift propellers 112, 212 may each have two blades 120, 220 that may be locked for cruising in minimum drag positions in which one blade is directly in front of the other blade as illustrated in
In some embodiments, the aircraft may include a single wing 104, 204 on each side of fuselage 102, 202 (or a single wing that extends across the entire aircraft). At least a portion of lift propellers 112, 212 may be located rearward of wings 104, 204 and at least a portion of tilt propellers 114, 214 may be located forward of wings 104, 204. In some embodiments, all of lift propellers 112, 212 may be located rearward of wings 104, 204 and all of tilt propellers 114, 214 may be located forward of wings 104, 204. According to some embodiments, all lift propellers 112, 212 and tilt propellers 114, 214 may be mounted to the wings—i.e., no lift propellers or tilt propellers may be mounted to the fuselage. In some embodiments, lift propellers 112, 212 may be all located rearwardly of wings 104, 204 and tilt propellers 114, 214 may be all located forward of wings 104, 204. According to some embodiments, all lift propellers 112, 212 and tilt propellers 114, 214 may be positioned inwardly of the ends of the wing 104, 204.
In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted to wings 104, 204 by booms 122, 222. Booms 122, 222 may be mounted beneath wings 104, 204, on top of the wings, and/or may be integrated into the wing profile. In some embodiments, lift propellers 112, 212 and tilt propellers 114, 214 may be mounted directly to wings 104, 204. In some embodiments, one lift propeller 112, 212 and one tilt propeller 114, 214 may be mounted to each boom 122, 222. Lift propeller 112, 212 may be mounted at a rear end of boom 122, 222 and tilt propeller 114, 214 may be mounted at a front end of boom 122, 222. In some embodiments, lift propeller 112, 212 may be mounted in a fixed position on boom 122, 222. In some embodiments, tilt propeller 114, 214 may mounted to a front end of boom 122, 222 via a hinge. Tilt propeller 114, 214 may be mounted to boom 122, 222 such that tilt propeller 114, 214 is aligned with the body of boom 122, 222 when in its cruise configuration, forming a continuous extension of the front end of boom 122, 222 that minimizes drag for forward flight.
In some embodiments, aircraft 100, 200 may include, e.g., one wing on each side of fuselage 102, 202 or a single wing that extends across the aircraft. According to some embodiments, the at least one wing 104, 204 is a high wing mounted to an upper side of fuselage 102, 202. According to some embodiments, the wings include control surfaces, such as flaps and/or ailerons. According to some embodiments, wings 104, 204 may have designed with a profile that reduces drag during forward flight. In some embodiments, the wing tip profile may be curved and/or tapered to minimize drag.
In some embodiments, rear stabilizers 106, 206 include control surfaces, such as one or more rudders, one or more elevators, and/or one or more combined rudder-elevators. The wing(s) may have any suitable design. In some embodiments, the wings have a tapering leading edge.
In some embodiments, lift propellers 112, 212 or tilt propellers 114, 214 may canted relative to at least one other lift propeller 112, 212 or tilt propeller 114, 214. As used herein, canting refers to a relative orientation of the rotational axis of the lift propeller/tilt propeller about a line that is parallel to the forward-rearward direction, analogous to the roll degree of freedom of the aircraft. Canting of the lift propellers and/or tilt propellers may help minimize damage from propeller burst by orienting a rotational plane of the lift propeller/tilt propeller discs (the blades plus the hub onto which the blades are mounted) so as to not intersect critical portions of the aircraft (such areas of the fuselage in which people may be positioned, critical flight control systems, batteries, adjacent propellers, etc.) or other propeller discs and may provide enhanced yaw control during flight.
As disclosed herein, the forward electric propulsion systems and aft electric propulsion systems may be of a clockwise (CW) type or counterclockwise (CCW) type. Some embodiments may include various forward electric propulsion systems possessing a mixture of both CW and CCW types. In some embodiments, the aft electric propulsion systems may possess a mixture of CW and CCW type systems among the aft electric propulsion systems.
Some embodiments may include an aircraft 400 possessing forward and aft electric propulsion systems where the amount of CW types 424 and CCW types 426 is not equal among the forward electric propulsion systems, among the aft electric propulsion systems, or among the forward and aft electric propulsion systems.
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 high voltage channels or power connection channels. Some embodiments may include various batteries being stored within an aircraft wing with high voltage channels traveling throughout the aircraft, including the wing and boom, to an electric propulsion system. In some embodiments, multiple high voltage power systems may be used to create an electric propulsion system with multiple high voltage power supplies to avoid the risk of a single point of failure. In some embodiments, an aircraft may include multiple electric propulsion systems that may be wired in a pattern to various batteries or power sources stored throughout the aircraft. It is recognized that such a configuration may be beneficial as to avoid the risk of a single point of failure where one battery or power source failure could lead to a portion of the aircraft not being able to maintain a required amount of thrust to continue flight or perform a controlled landing. For example, if a VTOL possessed two forward electric propulsion systems and two aft propulsion systems, the forward and the aft 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 were to fail, a forward and an aft electric propulsion system on opposite sides of the VTOL aircraft would remain in working order and may provide a more balanced flight or landing compared to a forward and aft electric propulsion system failing on the same side of a VTOL aircraft. Some embodiments may include four forward electric propulsion systems and four aft electric propulsion systems where diagonally opposing electric engines are connected to a common battery or power source. Some embodiments may include various configurations of electric engines electrically connected to high voltage power systems such that a risk of a single point of failure is avoided in the case of a power source failure and the phase of flight during which a failure occurs may continue or the aircraft may perform an alternative phase of flight in response to the failure.
As discussed above, an electric propulsion system may include an electric engine that provides mechanical shaft power to a propeller assembly to produce thrust. In some embodiments, the electric engine of an electric propulsion system may include a High Voltage Power System supplying high voltage power to the electric engines and/or a Low Voltage System supplying low voltage direct current power to an electric engine. Some embodiments may include the electric engine(s) digitally communicating with a Flight Control System (“FCS”) comprising Flight Control Computers (“FCC”) that may send and receive signals to and from the electric engine including commands and responsive data or status. Some embodiments may include an electric engine capable of receiving operating parameters from and communicating operating parameters to the FCC, including speed, voltage, current, torque, temperature, vibration, propeller position, and any other value of operating parameters.
In some embodiments, a flight control system may include a system capable of communicating with an electric engine to send and receive analog/discrete signals to the electric engine and controlling an apparatus capable of redirecting thrust of the tilt propellers between a primarily vertical direction during vertical flight mode to a mostly horizontal direction during forward-flight mode. In some embodiments, this system may be referred to as a Tilt Propeller System (“TPS”) and may be capable of communicating and orienting additional features of the electric propulsion system.
Some embodiments may include an electric propulsion system 602 including an electric engine subsystem 604 receiving signals from and sending signals to a flight control system 612. In some embodiments, a flight control system 612 may comprise a flight control computer capable of using Controller Area Network (“CAN”) data bus signals to send commands to the electric engine subsystem 604 and receive status and data from the electric engine subsystem 604. It should be understood that while CAN data bus signals are used between the flight control computer and the electric engine(s), some embodiments may include any form of communication with the ability to send and receive data from a flight control computer to an electric engine. In some embodiments, a flight control system 612 may also include a Tilt Propeller System (“TPS”) 614 capable of sending and receiving analog, discrete data to and from the electric engine subsystem 604 of the tilt propellers. A tilt propeller system 614 may include an apparatus capable of communicating operating parameters to an electric engine subsystem 604 and articulating an orientation of the propeller subsystem 606 to redirect the thrust of the tilt propellers during various phases of flight using mechanical means such as a gearbox assembly, linear actuators, and any other configuration of components to alter an orientation of the propeller subsystem 606.
As discussed throughout, an exemplary VTOL aircraft may possess various types of electric propulsion systems including tilt propellers and lift propellers, including forward electric engines with the ability to tilt during various phases of flight, and aft electric engines that remain in one orientation and may only be active during certain phases of flight (i.e., take off, landing, and hover).
In some embodiments, a motor and gearbox assembly 704 may contain a gearbox 710 aligned along the shaft 724 to provide a gear reduction between the torque of the shaft 724 from the electric engine assembly, comprising a stator 706 and rotor 708, and the output shaft 738 Torque applied to the output shaft 738 may be transferred to the propeller assembly 720. Some embodiments may include a gearbox 710 containing an oil pump. In such an embodiment, the oil pump may drive a circulation of oil throughout the motor and gearbox assembly 704 at a speed equivalent 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 a circulation of oil at a speed greater than or less than the rotation of the output shaft 738. Some embodiments of a motor and gearbox assembly 704 may include propeller position sensors 712 present within the housing that may detect a magnetic field produced by the electric engine assembly to determine a propeller position. Further embodiments may include propeller position sensors 712 that are powered by an inverter 716 and send collected data to an inverter 716.
In some embodiments, an electric engine assembly 702 may also include an inverter assembly 714 aligned along the shaft 724. An inverter assembly 714 may include an inverter 716 and an inverter power supply 740 An inverter power supply 740 may accept low voltage DC power from a low voltage system 734 located outside the electric engine assembly 702. An 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, that has been converted to low voltage DC power via a DC-DC converter 742. An inverter 716 may supply high voltage alternating current to the stator 706 of the electric engine assembly located within the motor and gearbox assembly 704 via at least one three-phase winding. An inverter assembly 714 may include an inverter 716 that may receive flight control data from a flight control computing subsystem 736.
In some embodiments, a motor and gearbox 704 may be located between an inverter assembly 714 and a propeller assembly 720. Some embodiments may also include a divider plate 744 coupled to the motor and gearbox assembly 704 and inverter assembly 714. A divider plate 744 may create an enclosed environment for an upper portion of the motor and gearbox assembly 704 via an end bell assembly, and create an enclosed environment for a lower portion of the inverter assembly 714 via a thermal plate. In some embodiments, divider plate 744 may serve as an integral mounting bracket for supporting heat exchanger 718. Heat exchanger 718 may comprise, for example, a folded fin or other type of heat exchanger. In some embodiments, the electric propulsion system 700 may circulate oil or other coolant throughout the electric engine assembly 702, motor and gearbox assembly 704, or inverter assembly 714 to transfer heat generated from the components to the oil or other coolant liquid. The heated oil or other coolant liquid may circulate through heat exchanger 718 to transfer the heat to an air flow 722 passing through the fins of the heat exchanger.
In some embodiments, the electric engine assembly 702 may be mounted or coupled to a boom structure 726 of the aircraft. A variable pitch mechanism 730 may be mechanically coupled to the propeller assembly 720. In some embodiments, the variable pitch mechanism may abut the electric engine assembly 702. In some embodiments, the variable pitch mechanism 730 may be coupled to the variable pitch mechanism 730 such it may be remotely mounted within the boom, wing, or fuselage of the aircraft. In some embodiments, the variable pitch mechanism 730 may include a shaft or component traveling within or adjacent to the shaft 724 to the propeller assembly 720. A variable pitch mechanism 730 may serve to change the collective angle of the forward electric engine's propeller assembly blades as needed for operation during the hover-phase, transition phase, and cruise-phase. Some embodiments may include the electric engine assembly 702 being mechanically coupled to a tilt propeller subsystem 728 that may redirect thrust between a primarily vertical direction during vertical flight mode to a mostly horizontal direction during forward-flight mode. In some embodiments, the tilt propeller subsystem may abut the variable pitch mechanism 730. Some embodiments may include a tilt propeller subsystem 728 comprising various components located in various locations. For example, a component of the tilt propeller subsystem may be coupled to the electric engine assembly 702 and other components may be coupled to the variable pitch mechanism 730. These various components of the tilt propeller subsystem 728 may work together to redirect the thrust of the tiltable electric propulsion system 700.
As discussed herein, a lift electric propulsion system may be configured to provide thrust in one direction and may not provide thrust during all phases of flight. For example, a lift system may provide thrust during take-off, landing, and hover, but may not provide thrust during cruise.
In some embodiments, a motor and gearbox assembly housing 904 may contain a gearbox 910 aligned along the shaft 940 to provide a gear reduction between the torque of the shaft 932 from the electric engine assembly, comprising a stator 906 and rotor 908, and the output shaft 932. Torque applied to the output shaft 932 may be transferred to the propeller assembly 920. Some embodiments may include a gearbox 910 containing a fluid pump for circulating cooling and/or lubrication fluid. In the embodiment shown, the fluid pump is an oil pump. In such an embodiment, the oil pump may drive a circulation of oil throughout the motor and gearbox assembly housing 904 at a speed equivalent to the rotation of the output shaft 932 to cool and lubricate the gearbox and electric motor components. Some embodiments of a motor and gearbox assembly housing 904 may include propeller position sensors 912 present within the housing that may detect a magnetic field produced by the electric engine assembly to determine a propeller position. Further embodiments may include propeller position sensors 912 that are powered by an inverter 916 and send collected data to an inverter 916 that may be transferred to a flight control computing system 930 among other flight control data.
In some embodiments, an electric engine assembly 902 may also include an inverter assembly housing 914 aligned along an axis sharing the axis of the shaft 924. An inverter assembly housing 914 may include an inverter 916 and an inverter power supply 934. An inverter power supply 934 may accept low voltage DC power from a low voltage system 928 located outside the electric engine assembly 902. An 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, that has been converted to low voltage DC power via a DC-DC converter 936. An inverter 916 may supply high voltage alternating current to the stator 906 of the electric engine assembly located within the motor and gearbox assembly housing 904 via at least one three-phase winding. An inverter assembly 914 may include an inverter 916 that may send data to and receive data from a flight control computing subsystem 930.
In some embodiments, a motor and gearbox housing 904 may be located between an inverter assembly housing 914 and a propeller assembly 920. Some embodiments may also include a divider plate 938 coupled to the motor and gearbox assembly housing 904 and inverter assembly housing 914. A divider plate 938 may create an enclosed environment for an upper portion of the motor and gearbox assembly housing 904 via an end bell assembly, and may create an enclosed environment for a lower portion of the inverter assembly housing 914 via a thermal plate. In some embodiments, a divider plate 938 may serve as an integral mounting bracket for supporting heat exchanger 918. Heat exchanger 918 may comprise, e.g., a folded fin or other type of heat exchanger. In some embodiments, the electric propulsion system 900 may circulate oil or other coolant fluid throughout the electric engine assembly 902, motor and gearbox assembly 904, or inverter assembly 914 to transfer heat generated from the components to the oil or other coolant liquid. The heated oil or other coolant liquid may be circulated through heat exchanger 918 to transfer the heat to an air flow 922 passing through the fins of the heat exchanger.
In some embodiments, a tiltable electric propulsion system and a lift electric propulsion system may possess similar components. This may be advantageous with respect to many design considerations present within VTOL aircrafts. For example, from a manufacturability standpoint, different types of electric propulsion systems having similar components may be beneficial in terms of manufacturing efficiency. Further, having similar components may be beneficial in terms of risk management as similar components possess similar points of failure and these points of failure may be well explored and designed around when comparing systems having similar components to systems having different components and configurations.
While a tiltable electric propulsion system may possess additional, and in some embodiments different, components compared to a lift electric propulsion system, it should be understood that in some embodiments a tiltable electric propulsion system and a lift electric propulsion system may possess the same configuration of components. For example, in some embodiments, a tiltable and lift electric propulsion system may contain the same components while the lift electric propulsion system may be coupled to a boom, wing, or fuselage of the aircraft such that it may not be able to provide thrust in as many directions as tiltable electric propulsion system.
Some embodiments of the disclosed electric engine may generate heat during operation and may comprise a heat management system to ensure components of the electric engine do not fail during operation. In some embodiments, coolant may be used and circulated throughout 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 assist with managing the heat present in the engine. Some embodiments may include using air cooling methods to cool the electric engine or using a mixture of coolant and air to manage the heat generated during operation in the electric engine. In some embodiments, the coolant being used may also be the same liquid that is being used as lubricant throughout the inverter, gearbox, or motor. For example, components of the electric engines may be cooled using a liquid or air or using a mixture of air and liquid cooling. As another example, a motor may be cooled using air cooling while the inverter and gearbox are cooled using liquid cooling. It should be understood that a mixture of cooling may be used for any combination of electric engine components or within each component.
In some embodiments, oil may be used as a lubricant throughout an electric engine and may also be used as coolant fluid to assist in managing the heat generated by the engine during operation. Further to this example, different amounts of oil may be used to act as both lubricant and coolant fluid in the electric engine, such as less than or equal to one quart, 1.5 quarts, two quarts, 2.5 quarts, three quarts, five quarts or any other amount of oil needed to lubricate and cool the electric engine, in combination with or without the assistance of air cooling. In some embodiments, the amount of the oil or liquid to be used in the system in relation to cooling may be determined based on an amount of thermal mass needed to drive heat transfer from the components of the electric propulsion system. As has been disclosed herein, an electric engine may have different primary functionalities such as being used only for lifting and landing, and as such only being used in one orientation, or being used during all stages of flight such as lifting, landing, and in-flight. An engine that is used in all stages of flight may experience various orientations throughout flight and may comprise more lubricant and coolant than the engine only used in one orientation. As such, all the engines on an aircraft may not include the same amount of lubricant and coolant. For example, a lifting and landing engine may only require less than one quart of oil while an engine that operates in all stages of flight may require more than one quart of oil. In some embodiments, the amount of oil or liquid for cooling may be of an appropriate amount to provide sufficient thermal mass to drive heat transfer from the components of the electric propulsion system no matter the orientation of the electric propulsion system. The embodiments discussed herein are exemplary, non-limiting, and do not dictate the bounds of the amount of lubricant and coolant that may be used in an electric engine.
Some embodiments may use oil to lubricate the electric engine and to cool the electric engine. Such embodiments may require additional volumes of oil. In such embodiments, the additional oil may allow for removal of traditional components that may be used to cool such an electric engine. For example, if the electric engine were cooled by another liquid such as glycol, the engine may comprise separate heat exchangers for both the lubricant fluid and the coolant fluid. As such, in embodiments where a single fluid is being used for both lubrication and cooling, such as oil, an increase in oil would be present but there would only be a need for one heat exchanger, so there may be a decrease in mass, due to using less heat exchangers and potentially other components not being required, of the overall system and a more appealing drag profile may be present. Further, using one substance for the lubrication and cooling of the engine may increase efficiency of the system due to the reduction in mass and the benefits of cooling the engine with a substance rather than relying on air cooling which may have issues traveling throughout the engine.
Some embodiments of electric engines may include various components for monitoring flammable fluids, and for preventing ingress of flammable materials into certain sections of the electric engine. Some embodiments may include an electric engine possessing a wet zone enclosure that may be defined by a gearbox, motor, and/or heat exchanger. In some embodiments, an electric engine may possess up to 4 liters, or more, of air within the motor-gearbox housing which is in contact with engine oil. Embodiments of a motor-gearbox housing may equalize internal and external pressure using a breather. Embodiments of a breather may include it protruding above nearby design features to prevent inadvertent entry of external fluids. Some embodiments may include a breather that possesses a screen and a circuitous entry path to prevent entry of external debris. Embodiments may include a sight glass being present on both the tilt and lift electric engines in order to check that oil is not overfilled or underfilled during servicing.
Some embodiments of electric engines may include active protection features in the forward and aft electric engines such as monitoring vibration throughout the engine and internal temperatures such as 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 hot-spot temperatures, and other various operating conditions throughout the engine as needed. Such monitoring may be accomplished using various sensors positioned throughout the electric propulsion system and aircraft. Embodiments may include vibration limits based on known failure points or resonances of components and overtemperature limits set based on known failure temperatures and operating limits in relation to auto-ignition temperatures of fluids. In some embodiments, the various sensors used to monitor the operating conditions throughout the engine may report operating conditions to the flight control system. Some embodiments may include a threshold operating value that may be required before an operating value is sent to, or flagged by, the flight control system. In some embodiments, a flight control system may, in response to detecting an operating condition, act to reduce the amount of power directed to an electric propulsion system. Some embodiments may include reducing the amount of power to an electric propulsion system to reduce mechanical wear or friction sparks from vibrations and/or reducing power in an effort to reduce the temperature of components present within the electric propulsion system. Further, some embodiments may include reducing power to an electric propulsion system where a detected efficiency of an inverter is less than a targeted efficiency. In some embodiments, for example where twelve electric propulsion systems are present within the aircraft, a flight control system may act to reduce power, or terminate power, to a single electric propulsion system while increasing the power directed to the remaining electric propulsion systems, or a subset thereof, to counter reduction in lift produced by the one electric propulsion system. In some embodiments, the flight control system may establish various thresholds of operating conditions to correspond with the reduction or increase of power to an electric propulsion system.
Some embodiments may include a High Voltage Power System that may have fuses at the high voltage battery terminals which may rapidly and irreversibly disconnect the engine electrical connection to mitigate and avoid overcurrent events. Such overcurrent protection may be activated when the electric engine current draw is greater than the Overcurrent operating. As such, in some embodiments, failure conditions which lead to overcurrent may only lead to a transient overheating, arc or spark faults. Some embodiments may include a fire threat characterization test ignition source that may be selected to be a more severe ignition source than a short occurring in the electric engine and being opened by the engine fuse. In some embodiments, an inverter may detect AC overcurrent and isolate the erroneous phase and/or will continuously monitor input DC voltage, and will apply protective actions to keep voltages under the overvoltage operating limit.
During takeoff, landing, hover and cruise, motors and related control components of the VTOL aircraft may generate heat. The heat must be dissipated to prevent degradation or damage to the motor, control components and other elements of the VTOL aircraft. For some types of VTOL aircraft, such as electric VTOL (eVTOL) aircraft, thermal control is likewise important to maintain optimal energy efficiency of, e.g., battery-powered components.
Some elements may generate high thermal loads only during certain operational periods. For example, some lift propellers may be used only during takeoff, landing, and hover, and may be shut off during cruise. Therefore, such lift propellers may generate a high thermal load during takeoff, landing, and hover, and generate little or no heat during cruise.
As described herein, embodiments of an electric engine may include an inverter assembly, a gearbox assembly, and an electric motor assembly, or various combinations thereof. In some embodiments, the inverter assembly, the gearbox assembly, and the electric motor assembly may be substantially aligned along a central axis of the electric engine. As disclosed herein, these assemblies or combinations thereof may be substantially aligned along an axis by sharing a common axis or having parallel axes that are within a distance less than or equal to 5% of the outer diameter of the component with the largest diameters of one another. For example, the inverter assembly, the gearbox assembly, and the electric motor assembly may be substantially aligned along a central axis where the central axis of the inverter assembly, the central axis of the gearbox assembly, and the central axis of the electric motor assembly are within a distance less than or equal 5% of the outer diameter of the electric motor assembly where the electric motor assembly possesses a greater outer diameter than the gearbox assembly and inverter assembly. It should be understood that the embodiments described herein are only exemplary and while certain components of the electric propulsion system may be shown to abut others, all configurations of abutting may be present. For example, a gearbox assembly may be shown to abut an inverter assembly and an electric motor assembly. Further, in some embodiments an inverter assembly may abut a gearbox assembly and an electric motor assembly. Some embodiments may include an electric motor assembly abutting a gearbox assembly and an inverter assembly.
In some embodiments, each of the inverter assembly, the gearbox assembly, and the electric motor assembly may abutting at least one of the other assemblies. Abutting may include direct or indirect contact between the components comprising an assembly or housings wherein an assembly is located. In some embodiments, an electric engine may include an inverter assembly and an electric motor assembly without a gearbox assembly. Some embodiments of an electric engine may include an electric motor assembly and a gearbox assembly without an inverter assembly, or an electric motor assembly without a gearbox assembly or an inverter assembly.
In some embodiments, an electric engine may include a gearbox.
In some embodiments, an electric engine 1100B may include bearings 1124B, 1126B aligned along the main shaft 1110B. Some embodiments may include an inner race of bearings 1124B, 1126B that are mechanically coupled to the planetary carrier and various bearings such as 1124B and 1126B.
Electric propulsion systems 1100A-C, as discussed above, are exemplary embodiments. However, it is understood that while electric propulsion system 1100A may be capable of providing required thrust to a VTOL aircraft, it may creating a larger drag profile and contribute more mass to the VTOL aircraft than electric propulsion systems 1100B and 1100C. Electric propulsion systems 1100B and 1100C comprise gearbox assemblies. As such, the electric propulsion systems 1100B and 1100C possess a gear reduction that allows the electric motor assembly, and thus the electric propulsion systems, to possess smaller drag profiles and less mass.
Electric propulsion system 1100B may possess a gearbox assembly between an electric motor assembly and a shaft flange assembly. While this configuration may require less mass than electric propulsion system 1100A, it may require more mass than electric propulsion system 1100C. Electric propulsion system 1100B may possess a gearbox assembly such that an input shaft, or sun gear, travels from the electric motor assembly to the gearbox assembly and an output shaft, or portion of planetary carrier, travels from the gearbox assembly to the shaft flange assembly. Whereas electric propulsion system 1100C, in some embodiments, may possess a sun gear that travels from the rotor of the electric motor assembly to the gearbox assembly and a main shaft, coupled to a planetary carrier or carrier cover, that travels through the sun gear, past the electric motor assembly to a shaft flange assembly. As such, electric propulsion system 1100C may comprise a more compact design, housing, and drag profile when compared to electric propulsion system 1100B. This may result in a more efficient drag profile and a more mass efficient system. Further, electric propulsion system 1100B may possess a gearbox assembly without means of lubrication which may limit the run time of the electric propulsion system. Electric propulsion system 1100C may comprise a heat exchanger to cool and lubricate portions of the systems, including the gearbox assembly. This may lead to additional efficiency and long flight range times.
Due to the mechanical coupling as described herein, torque may be transferred from the rotor 4904 to sun gear 4906 along path 4930. The sun gear 4906 may transfer torque to planetary gear 4908 along paths 4932 and 4934. The planetary gear 4908, and in some embodiments compound planetary gear 4908, 4950, may transfer torque to planetary carrier along paths 4936 and 4938. The planetary carrier may transfer torque to the main shaft 4918 along paths 4940. The main shaft 4918 may transfer torque along its length, via path 4944 and to a propeller assembly 4920. A propeller assembly 4920 may transfer torque to propellers via paths 4946 and 4948. It should be understood that the paths discussed above are exemplary and all configurations consisting of sending torque away from a propeller to a gearbox assembly and then sending the torque back through the gearbox assembly and electric motor assembly to the propellers are considered.
In some embodiments, a process of delivering power from an electric engine using a gearbox assembly via a reverse torque path may comprise driving a planetary gear that is mechanically coupled to a rotor of an electric motor assembly. In some embodiments, the planetary gear may interface with a sun gear and a ring gear. Some embodiments may include a hollow sun gear and a fixed ring gear. Some embodiments may include driving a planetary carrier that is connected a shaft that extends from the planetary gear. A shaft that extends from a planetary gear may include a shaft aligned along a central axis of the planetary gear. Some embodiments may include driving a carrier cover that is connected to a shaft from a shaft that extends concentrically from the planetary gear. Some embodiments may include driving a main shaft. Driving a main shaft may comprise driving a first portion of the main shaft that is mechanically coupled to the carrier cover and transferring torque along the main shaft to a second portion of the main shaft that is mechanically coupled to a propeller assembly. Some embodiments may include a heat exchanger, as described herein, that may cool the gearbox assembly using various amounts of oil that may comprise one quart, 1.5 quarts, two quarts, 2.5 quarts, three quarts, or five quarts.
In some embodiments, a gearbox assembly may include multiple sets of gearboxes. For example, in some embodiments, the output of a gearbox assembly may be fed into another gearbox assembly to achieve a greater gear reduction. Such embodiments may include at least one sun gear, at least one set of planetary gears, at least one ring gear, and at least one planetary carrier. The gearboxes may possess common gears such as a common sun gear, a common set of planetary gears, and a common ring gear. The embodiments discussed herein may be modified to include multiple sets of gearboxes.
At block 5304, a process for delivering power from an electric engine using a gearbox assembly may include driving a planetary carrier that is connected to at least one shaft that extends concentrically from the planetary gear, consistent with the discussion throughout this disclosure.
At block 5306, a process for delivering power from an electric engine using a gearbox assembly may include driving a carrier cover that is connected to at least one shaft from a set of shafts that extends concentrically from the planetary gear, consistent with the discussion throughout this disclosure.
At block 5308, a process for delivering power from an electric engine using a gearbox assembly may include driving a main shaft, consistent with the discussion throughout this disclosure.
At block 5310, a process for delivering power from an electric engine using a gearbox assembly may include driving a first portion of the main shaft that is mechanically coupled to the carrier cover, consistent with the discussion throughout this disclosure.
At block 5312, a process for delivering power from an electric engine using a gearbox assembly may include transferring torque along the main shaft to a second portion of the main shaft that is mechanically coupled to a propeller assembly, consistent with the discussion throughout this disclosure.
As described herein, an electric propulsion system may include an inverter assembly, gearbox assembly, and engine assembly. In some embodiments, the electric propulsion system 1200A may include components packaged in various housings including a motor-gearbox assembly housing 1202A and an inverter assembly housing 1228A. Enclosing the various components of electric propulsion system 1200A in housings 1202A and 1228A may provide various benefits, including lower mass and a more efficient drag profile, as described herein. Further, in some embodiments, the gearbox assembly, inverter assembly, and/or electric motor assembly may possess a substantially circular profile. As used herein a profile may be substantially circular where the length of a minor axis of a circular shape and the length of a major axis of a circular shape possess a relationship such that the length of the minor axis is at least a threshold amount, such as 80%, of the length of the major axis. Further, in some embodiments, the gearbox assembly, inverter assembly, and electric motor assembly, or a subset of those listed, may be sized such that the assemblies possess substantially equivalent radii. As used herein assemblies may possess substantially equivalent radii where the difference among the radii between two assemblies is less than a threshold amount, such as 10%, of the radius of the largest assembly. In some embodiments, the profile of components making up the electric propulsion system as described herein may include various polygons such as hexagons, heptagons, octagons, nonagons, decagons, and additional polygons have more than ten sides.
Electric propulsion system 1200A may include an electric motor assembly, including stator 1204A, rotor magnet 1206A, and rotor 1208A.
In some embodiments, the electric motor assembly may interact with, and in some embodiments transfer toque to, a gearbox assembly. Electric propulsion system 1200A may include a gearbox assembly, comprising a sun gear 1214A, a set of planetary gears 1216A, a planetary carrier 1218A, and a carrier cover 1220A. Some embodiments may include a sun gear 1214A having teeth that interact with teeth of the planetary gears 1216A, and a ring gear (not picture here in this figure) having teeth that also interact with the teeth of the planetary gears 1216A. In some embodiments, a shaft 1222A may extend through or from the planetary gears 1216A. In some embodiments, the planetary carrier 1218A may receive a first end of the shafts 1222A such that the planetary carrier 1218A may rotate at the same rate as the planetary gears 1216A. In some embodiments, the carrier cover 1220A may receive a second end of the shafts 1222A such that the carrier cover 1220A may rotate at the same rate as the planetary gears 1216A. In some embodiments, the planetary gears 1216A, planetary carrier 1218A, and carrier cover 1220A may be mechanically coupled along the axis of shaft 1222A.
In some embodiments, electric propulsion system 1200A may include a main shaft 1210A that may be mechanically coupled to a shaft flange assembly 1224A to provide mechanical shaft power to turn the propellers of a propeller assembly. As used herein, components may be mechanically coupled where there exists any connections or coupling, whether direct or indirect, between two components. A shaft flange assembly may include a flange that is coupled to a main shaft with a splined connection to take torque loads from the main shaft and transfer the torque to the propellers that coupled to the flange. A flange may also be coupled to a main shaft using fasteners, by welding, by brazing, or any other use of components or methods to couple the main shaft and the flange. In some embodiments, a main shaft and a flange may be machined together to form a single component. In some embodiments, a shaft flange assembly may be a component of a propeller assembly that may comprise a shaft flange assembly, propellers, and a spinner. In some embodiments, a shaft flange assembly may also be referred to as a propeller hub.
In some embodiments, electric propulsion system 1200A may include components for an inverter assembly, as described herein. For example, electric propulsion system 1200A may include printed circuit board assemblies (PCBAs) such as a power PCBA 1230A that may comprise power modules 1232A, a gate drive PCBA 1236A, and a control PCBA 1240A, Some embodiments of the inverter assembly of the electric propulsion system 1200A may also include a spacer board 1238A among the various PCBAs. Further, some embodiments may include an energy storage device, for example a DC capacitor that may be stored within the DC capacitor housing 1234A. Some embodiments of an inverter assembly may also include busbar connectors 1244A to supply alternating current to the electric motor assembly. Some embodiments of an inverter assembly may include power connections 1246A coupled to a high voltage connector to deliver high voltage power to the inverter assembly.
Some embodiments of an inverter assembly of the electric propulsion system 1200A may include layering the respective inverter assembly components in a stacking formation along guide pins 1242A that extend through each layer of the inverter assembly. It is recognized that an inverter assembly utilizing a stacking formation along guide pins 1242A may be beneficial in various design criteria relevant for VTOL aircrafts. For example, a stacking formation may allow for a more compact packaging of the inverter assembly, and thus may help in minimizing the mass of the electric propulsion system 1200A and minimize the drag experienced due to the electric propulsion system packaging. Further, a stacking formation of the inverter assembly may be advantageous from a manufacturing perspective as a stacking formation may allow for tolerances within various parts of the inverter assembly. In some embodiments, structural components may be introduced to the inverter assembly to assist in supporting the stacking formation with loads experienced during various phases of flight. Some embodiments may include inverter assembly components also acting as structural components. For example, a DC capacitor housing 1234A may house the capacitor, as well as other components, for the inverter assembly and may be made of a plastic, or other material, capable of supporting the PCBAs and other components surrounding it.
In some embodiments, an electric propulsion system 1200A may include a heat exchanger 1226A coupled to the motor-gearbox assembly housing 1202A and an inverter assembly housing 1228A. A heat exchanger 1226 may be coupled to a dividing plate comprising a thermal plate 1248A and an end bell plate 1250A. An end bell plate 1250A may serve to close off the motor-gearbox assembly housing 1202A. A thermal plate 1248A may serve to close off the inverter assembly housing 1228A. In some embodiments, a dividing plate may serve as an integral mounting bracket for supporting heat exchanger 1226A. Heat exchanger 1226A may comprise, e.g., a folded fin or other type of heat exchanger. In some embodiments, the electric propulsion system 1200A may circulate oil or other coolant throughout the electric motor assembly, gearbox assembly, or inverter assembly to transfer heat generated from the components to the oil or other coolant liquid. The heated oil or other coolant liquid may be circulated through the fins of heat exchanger 1226A by an internal liquid flow paths which may possess an inlet and outlet for the liquid flow paths that may be coupled to an outlet and inlet, respectively, of the bores or grooves that may be present on the dividing plate. In some embodiments, a motor-gearbox housing 1202A may comprise a sump 1212A. A sump 1212 may serve to collect oil or liquid coolant distributed throughout the electric propulsion system 1200A and recirculate the oil or liquid coolant.
In some embodiments, a heat exchanger may be fluidically coupled to the gearbox assembly, inverter assembly, and/or electric motor assembly. As used herein, an assembly, or components therein, may be fluidically coupled where a liquid flow path from the heat exchanger may interact with, supply liquid to, or interface with the assembly or components therein.
Embodiments of an electric engine may include an electric motor assembly, as described herein.
Disclosed embodiments of an electric motor assembly may include a rotor. In some embodiments, the electromagnetic field produced by a stator in an electric motor assembly may drive a rotation of a rotor about an axis.
As described herein, disclosed embodiments of an electric propulsion system may include a motor assembly and gearbox assembly. In some embodiments, a gearbox assembly may comprise torque paths that may exert loads. As described herein, an electric propulsion system may include a gearbox assembly and a rotor of an electric motor assembly, both of which may exert loads on a shaft. For example, gyroscopic effects due to the spinning rotor being in motion may exert moment loads. The moment loads may be on a centralized path of the shaft. A gearbox assembly, which may include planetary gears, may share torque through several paths, so the sharing of loads may be dependent on tolerances of components within the electric engine. As such, solutions that support loads and resist moments created by generated torque while maintaining a low mass and drag profile may be advantageous.
Disclosed embodiments may include a bearing system comprising a rotor utilizing a bearing to support loads.
Disclosed embodiments of an electric propulsion system may also include a pilot system for bearings to support a rotor. As disclosed herein, an electric propulsion system may include a bearing that supports a sun gear and rotor. A bearing supporting a rotor may comprise a bearing with an outer race mechanically coupled to an inner surface of a rotor. For example, bearing 1634C may comprise outer race mechanically coupled to rotor hub 1604C. Bearing 1634C may pilot sun gear 1612C and rotor hub 1604C, by guiding an alignment or mating of multiple components. For example, a pilot may serve to align or mate the sun gear 1612C and rotor hub 1604C. Bearing 1634C may support an edge of sun gear 1612C and an edge of rotor hub 1604C to rest on outer race, which may concentrically affix sun gear 1612C and rotor hub 1604C. A first edge of sun gear 1612C and a first edge of rotor hub 1604C may abut and meet on the outer race of bearing 1634C. Bearing 1634C may influence the diameter of sun gear 1612C and rotor. The diameter of outer race of bearing 1634C may be substantially similar to a diameter of an inner surface of sun gear 1612C and an inner diameter of a rotor. In some embodiments, rotor hub 1604C and sun gear 1612C may be concentrically affixed. Sun gear 1612C may have a diameter equal to an inner diameter of a rotor. In some embodiments, a pilot system for bearings may include shoulders. A shoulder may be an edge of a component that abuts one or more edges of another component. For example, shoulders may comprise a portion of the sun gear 1640B that abuts one or more edges of bearing 1616B, and a portion of the rotor hub 1642B that abuts one or more edges of bearing 1616B. Shoulders may cooperate to restrict movement of a bearing. For example, shoulders may cooperate to restrict movement of bearing 1616B in an axial direction of the shaft or along the axis 1624B. In some embodiments, a pilot may include shoulders to capture a bearing radially. A pilot system may reduce mass and prevent the need for additional materials. In some embodiments, dowel pins may be used to pilot a sun gear and rotor.
In some embodiments, a rotor bearing system may also include bearings to resist moment loads and allow float to compensate for tolerances in a gearbox. A rotor bearing system may include a hydrodynamic bearing. A hydrodynamic bearing may resist, or counteract, rotor moment loads. In some embodiments, a hydrodynamic bearing may be positioned along a sun gear. For example, a hydrodynamic bearing may be located between a sun gear 1612C and main shaft 1626C, and the hydrodynamic bearing may be located in a position along the length of the sun gear 1612C. In some embodiments, the hydrodynamic bearing may extend along the full length of the sun gear 1612C. The hydrodynamic bearing may be positioned where a main shaft 1626C has a shoulder, or cavity, as described herein. For example, the hydrodynamic bearing may comprise fluids between a sun gear 1612C and a shoulder, or cavity, of a main shaft 1626C. In some embodiments, the size or shape of the shoulder may be determined by properties of the rotor. For example, the shoulder may have a depth and width which may be determined by properties of the rotor including mass, speed, rate of change, and change in axis or loads (including gyroscopic, axial, and radial loads or moments). The hydrodynamic bearing may comprise fluids, such as oil, located between a sun gear 1612C and the outer surface 1628C of a main shaft 1626C. The hydrodynamic bearing may assist in resisting moment loads experienced by sun gear 1612C. For example, the hydrodynamic bearing may exert a restoring force to resist gyroscopic loads. In some embodiments, hydrodynamic bearing may comprise oil. The hydrodynamic bearing may allow sun gear 1612C or a ring gear to float. The hydrodynamic bearing may allow for tolerances within various components of the electric propulsion system. The hydrodynamic bearing may comprise the same liquid, such as oil, that is used throughout the electric propulsion system for lubrication and cooling. As discussed herein, utilizing a single liquid for hydrodynamic bearings, cooling, and lubricating may provide advantages of reducing mass and reducing the size of various components.
In some embodiments, the manufacturing constraints of the various areas of the rotor 5000 may determine the mass of the layer 5004, 5012. For example, layer 5004, 5012 may have a certain minimum or maximum mass that is dictated by manufacturing equipment and steps of creating the rotor. Layer 5004, 5012 may include a mass of rotor material that is sized with a magnitude capable of countering the manufacturing constraints present in each area of the rotor. In some embodiments, the mass of the layer 5004, 5012, to later be removed in balancing the rotor, may be determined based on the precision of the machine(s) manufacturing the various portions of the rotor. For example, in some embodiments a machine or machines may manufacture portions of the rotor with a precision of +/−5% of the target mass of the portion of the rotor. In such an example, the rotor may include a layer 5004, 5012 with an overall mass that includes enough mass to be removed in balancing a rotor with a mass falling within the compounded deviated mass due to the precision of the machine(s) manufacturing the rotor. In some embodiments, the material properties of the layer 5004, 5012 may include aluminum, steel, or another other material capable of accommodating the manufacturing precision of the machine(s) manufacturing the portions of the rotor. In some embodiments, the thickness, or width, and depth of layer 5004, 5012 may be thicker or thinner depending on design considerations, the needs of the system, and manufacturing constraints. In some embodiments, various layers may have substantially similar widths and depths, where substantially similar comprises difference in the widths or depths of the layers being less than 5% of the greater widths or depths.
In some embodiments, a rotor 5000 may include multiple layers 5004, 5012. For example, a rotor 5000 may include a layer 5004 on the edges of an inner surface, or circumference, of the rotor hub 5002, as shown in
In some embodiments, a layer 5004, 5012 may include grooves 5006 creating portions 5008. In some embodiments, the portions may be made of aluminum. Further, some embodiments may include grooves 5006 that may be made of aluminum. In some embodiments, grooves 5006 may serve to act as a liquid flow path for oil or other liquid that is present within the electric motor assembly. For example, in normal operation, as described herein, oil or liquid may be circulated throughout an electric motor assembly to assist in cooling or lubricating components. As such, grooves 5006 may act to allow oil or liquid to pass through the layers 5004, 5012 so that oil or liquid does not gather within the layers 5004, 5012 and is returned to the sump or other reservoir as described herein. In some embodiments, the multiple layers 5004 may be aligned such that the grooves 5006 of each layer are aligned.
In some embodiments, a rotor 5000 may comprise through-holes 5010. In some embodiments, a rotor 5000 may be machined with through-holes 5010. Those of ordinary skill in the art will appreciate that mass is a critical factor in aircraft design, and particularly in VTOL aircrafts design. Mass may impact the efficiency, payload, and flight time of a VTOL aircraft. As such, some embodiments of a rotor may include through-holes 5010, produced by a machining process, lasering process, or any other process of removing mass from a rotor. Through-holes 5010 may reduce the mass of a rotor 5000 by removing sections of the rotor hub 5002 material, such as aluminum. In some embodiments, through-holes may also serve as connection points for a sun gear as discussed above.
In some embodiments, an electric motor assembly of a VTOL aircraft, as described herein, may generate torque by rotating a rotor 5000 at high rates of rotation. At high rotational speeds, an unbalanced rotor having an axis of rotation that does not align with the center of mass of the rotor will experience high levels of unwanted vibration and noise. An unbalanced rotor may result from production tolerances of the manufacturing processes. For example, magnets present in the rotor assembly may not have a uniform mass and may not be uniformly placed along the rotor or lamination stack. Further, in some embodiments, through-holes 5010 may be produced using machining process and production tolerances may lead to an unbalanced rotor. As such, it is recognized that a process for balancing a rotor may be advantageous. It is also recognized that mass may be a critical design criteria in VTOL aircrafts and as such, traditional rotor balancing techniques that involve adding mass to the rotor or removing a minimal amount of mass may lead to unwanted mass remaining on the rotor. In some embodiments, processes may be used to balance a rotor while achieving a maximum reduction in the mass of the rotor.
Some disclosed embodiments may comprise an improved process for balancing a rotor of an electric motor assembly. Some embodiments may include identifying an axis of rotation of a rotor. As discussed herein, a rotor 5000 may include a layer 5004, 5012. In some embodiments, the layers 5004, 5012 serve as a sacrificial layer that may be machined to be integral with the rotor to later be removed from the rotor 5000 to achieve a balanced rotor. In some embodiments, the portions 5008 may serve as sacrificial portions that may be machined with the rotor to later be removed from the rotor 5000 to achieve a balanced rotor. It is recognized that removable layers and portions added to the rotor after the rotor is manufactured, via fasteners, glue, or similar materials, could serve a similar purpose as a sacrificial layer or sacrificial portion. However, removable layers and portions would require additional mass in the form of attachments that would not be advantageous to a VTOL's overall efficiency. Further, the attachments used for removable layers and portions, such as fasteners and adhesives, are at risk of failing during flight and may damage other components of the electric propulsion system.
Some embodiments of a process for balancing a rotor may comprise determining an imbalance present in the rotor by rotating the rotor about the axis of rotation. Determining an imbalance may include rotating the rotor and detecting a phase, and respective magnitude, of the imbalance. Some embodiments may include marking the rotor by laser etching the rotor, placing a reflective sticker on the rotor, or any other method of creating a distinctive mark on the rotor. In some embodiments, rotating the rotor may include using a machine to spin the rotor about the axis of rotation at a speed less than operating speed, such as a dynamic balancer. Operating speed may include the expected rate of rotation of the rotor for any phase of flight. In some embodiments, rotating the rotor may include rotating the rotor at a speed less than the first resonance of the rotor. Detecting a phase of an imbalance may include using a machine to monitor the distinctive mark of the rotor while in rotation. In some embodiments, the machine to monitor the rotation of the rotor may be the same machine that may rotate the rotor. The machine may be able to track the distinctive mark and calculate a displacement of the mark during rotation, indicating an imbalanced rotor. In some embodiments, detecting a phase of an imbalance may also include receiving signals from an encoder or accelerometer during the rotation, or downloading after the rotation, to identify the position of the rotor or forces experienced by the rotor in the position where the encoder, accelerometer, or a similar sensor is positioned on the rotor.
Some embodiments of a process for balancing a rotor may comprise calculating an amount of mass to add or remove at a position along the layer 5004, 5012 to correct an imbalance present in the rotor. Some embodiments may include a machine or algorithm that analyzes the phase, and respective magnitude, of the imbalance to determine amount of mass to add or remove and position for the mass to be added to or removed from. In some embodiments, a machine that calculates the mass to be added or removed may include the machine that is rotating the rotor, detecting the imbalance, or may be a separate machine. In some embodiments, the mass to be added or removed at a position along the layer 5004, 5012 may alter the center of mass of the rotor such that it coincides with the axis of rotation of the rotor. In some embodiments, a rotor may possess multiple layers 5004, 5012 posited along an inner surface of the rotor a distance from the edge of the rotor. As such, balancing the rotor may include balancing the rotor among one or more planes of the rotor by adding or removing mass along one, or along more than one, of the layers.
In some embodiments, removing an amount of mass from the layer 5004, 5012 may include machining away a portion of the volume of the layers 5004, 5012. In some embodiments, removing an amount of mass from the layers 5004, 5012 may include removing anywhere from 50% to 100% of the volume of the layers 5004, 5012. Removing 50% to 100% of the volume of the layers 5004, 5012 may reduce the mass of the rotor. In some embodiments, layers 5004, 5012 may only be present to be sacrificial material in balancing the rotor. Layers 5004, 5012 may be integrally formed with the rotor, to provide integrated rotor balancing material that is removed, rather than added. By removing sacrificial rotor material, a balancing process would not require the use of adhesives or fastening methods to add balancing weight.
In some embodiments, after removing material from the layers 5004, 5012, an amount of mass of the layers remaining on the rotor may be the minimal amount of mass required to balance the rotor, and thus may result in a balanced rotor with a minimized mass. Removing majority of the volume of the layers present on the rotor may allow for a reduction in mass of the rotor such that the rotor contains no additional material. For example, if a rotor with layers 5004, 5012 was determined to be balanced without removing any portion of the layers 5004, 5012, 100% of the volume comprising the layers 5004, 5012 may be machined away as none of the mass from the layers would be needed to balance the rotor. In some embodiments, it may be determined that 3% of the volume of layer 5004 would need to be present to balance in rotor. In such an example, 100% of layer 5012 may be removed and 97% of layer 5004 may be removed to balance the rotor.
Some embodiments may include utilizing specific machinery to remove the volume of layers 5004, 5012 in balancing the rotor. Some embodiments may include utilizing machinery capable of the volumes of layers 5004, 5012 at a precision of 0.01% to 0.1% of the layer. In some embodiments, machinery may be used in machining away the volume of the layers, such as a lathe or a CNC machine, at a resolution less than five microns. In such embodiments, using machinery capable of such precision may achieve the advantages of a balanced rotor having minimal mass. Some embodiments may include utilizing various types of machinery when removing the layers, such as removing a large fraction of the mass to be removed with a method with less precision than the method to remove the remaining amount of the mass to be removed.
Some embodiments may include calculating an amount of mass to be added at a position along the layers 5004, 5012 to balance to the rotor, and balancing the rotor may include machining away the volume of the layers 5004, 5012 such that only an amount of mass of the layers remaining is a portion of the layers 5004, 5012 that is equal to the amount of mass that was calculated to be added and present at the calculated position. In some embodiments, calculating an amount of mass to be removed at a position along the layers 5004, 5012 to balance the rotor may include machining away the volume of layers 5004, 5012 such that only an amount go mass of the layers remaining is a portion of the layers 5004, 5012 that is equal to the amount of mass that was calculated to be removed and present at a position on the opposite side of the layer from the calculated position.
In some embodiments, calculating an amount of mass to be removed may include calculating a number of sacrificial portions 5008 to remove. In some embodiments, the portions 5008 may be defined by grooves 5006. Some embodiments may include removing full portions 5008 or any partial amounts of portions 5008. In some embodiments, removing portions 5008 may include calculating a maximum amount of portions to be removed to achieve a balanced rotor. In some embodiments, the number k of portions to be removed may include an amount of portions such that if any additional mass were to be removed from the rotor after k portions are removed, the rotor may never be able to achieve balance.
At block 5104, a process for balancing a rotor of an electric engine of an electric propulsion system may include determining an imbalance present in the rotor by rotating the root about the axis of rotation, consistent with the discussion throughout this disclosure.
At block 5106, a process for balancing a rotor of an electric engine of an electric propulsion system may include calculating an amount of mass k to add at a position p along the sacrificial layer such that the center of mass of the rotor coincides with the axis of rotation of the rotor, consistent with the discussion throughout this disclosure.
At block 5108, a process for balancing a rotor of an electric engine of an electric propulsion system may include removing an amount of mass r from the sacrificial layer such that an amount of remainder mass n is present along the circumference of the rotor, consistent with the discussion throughout this disclosure.
As discussed herein, a rotor assembly of an electric motor assembly may comprise a rotor mechanically coupled to a sun gear. Similar to the discussion above with respect to balancing a rotor, it may be advantageous to balance the rotor assembly to avoid unwanted vibrations and noise during normal operation. A rotor assembly may be unbalanced due to manufacturing tolerances and due to the multiple mating parts throughout the rotor assembly.
In some embodiments, a process for balancing the rotor assembly may include identifying an axis of rotation of the rotor assembly and rotating the rotor assembly at speeds less than operating speed. In some embodiments, the rotor assembly may be coupled to a machine that may be capable of rotating the rotor assembly at speeds less than operating speed. In some embodiments the rotor assembly may be rotated at a speed less than the first resonance of the rotor assembly. Some embodiments may include determining an imbalance present in the rotor assembly. Determining an imbalance present in the rotor assembly may include using a machine to identify a phase, and magnitude, of the imbalance by tracking a distinctive mark on the rotor assembly, such as a reflective sticker or laser etched mark, or using an electric eye, encoder, accelerometer, or a similar component to track the motion of the rotor assembly.
Some embodiments may include calculating an amount of mass to add to the rotor assembly such that the center of mass of the rotor assembly coincides with the axis of rotation of the rotor assembly. Calculating an amount of mass, and its respective position, may be done using a machine or algorithm that analyzes the phase and magnitude of the imbalance in various planes of the rotor assembly to determine an amount of, and position of, mass to be added to the rotor assembly. In some embodiments, the mass to be added may be in the form of rivets. Rivets may include masses that may be removably attached or permanently affixed to the through-holes 5010 of rotor 5000. Rivets may be made of aluminum, copper, steel, or any other material capable of balancing the rotor assembly. Adding rivets may include permanently affixing or removably attaching rivets to the rotor via through-holes such that the rotor assembly is balanced. In some embodiments, the amount of mass to be added may include rivets possessing different material properties and positions.
As shown in
At block 5204, a process for balancing a rotor assembly may proceed to determining an imbalance present in the rotor by rotating the rotor about the axis of rotation, consistent with the discussion throughout this disclosure.
At block 5206, a process for balancing a rotor assembly of an electric engine of an electric propulsion system may include calculating an amount of mass k to add at a position p along the sacrificial layer such that the center of mass of the rotor coincides with the axis of rotation of the rotor, consistent with the discussion throughout this disclosure.
At block 5208, a process for balancing a rotor assembly may proceed to include removing an amount of mass r from the sacrificial layer such that an amount of remainder mass n is present along the circumference of the rotor, consistent with the discussion throughout this disclosure.
At block 5210, a process for balancing a rotor assembly may proceed to identifying an axis of rotation of a rotor assembly, wherein the rotor assembly comprises the rotor mechanically coupled to a sun gear, consistent with the discussion throughout this disclosure.
At block 5212, a process for balancing a rotor assembly may proceed to determining an imbalance present in the rotor assembly by rotating the rotor about the axis of rotation, consistent with the discussion throughout this disclosure.
At block 5214, a process for balancing a rotor assembly may proceed to calculating a number j of rivets to add to the rotor assembly such that the center of mass of the rotor assembly coincides with the axis of rotation of the rotor assembly, consistent with the discussion throughout this disclosure.
At block 5216, a process for balancing a rotor assembly of an electric engine of an electric propulsion system may include adding the j rivets to the rotor assembly, consistent with the discussion throughout this disclosure.
Disclosed embodiments of an electric propulsion system may include a gearbox assembly, as described herein. A gearbox may assist in a gear reduction for an electric propulsion system. Some embodiments of an electric propulsion system may include a gearbox assembly located between the electric motor assembly and the end bell assembly, as described herein. A gearbox assembly may comprise a main shaft assembly.
In some embodiments, a main shaft assembly 1700 may comprise a planetary carrier 1712 having bearings 1722 to assist the planetary carrier 1712 in receiving shafts 1710. Bearings 1722 may allow the shafts 1710 to rotate with the planetary gears 1704, 1706 while allowing the shafts to be housed within the planetary carrier 1712. In some embodiments, a carrier cover 1714 may have bearings 1718 to assist the carrier cover 1714 in receiving shafts 1708 such that the shafts 1708 may rotate with planetary gears 1704, 1706 while allowing the shafts 1708 to be housed within the carrier cover 1714. In some embodiments, washers 1720, 1724 may be positioned between the planetary gears 1704, 1706 and the carrier cover 1714 and planetary carrier 1712, respectively. Washers 1720, 1724 may be designed to account for machine tolerances in the manufacture of components throughout the gearbox assembly, or provide the planetary gears 1704, 1706 with a surface to rotate against without damaging the planetary carrier 1712 or carrier cover 1714. Some embodiments may include mechanically coupling the planetary carrier 1712 and carrier cover 1714 using screws 1728 or similar components.
In some embodiments, a carrier cover 1712 may be mechanically coupled to the main shaft 1702. In such an embodiment, a rotation of the main shaft would rotate at the same speed as the carrier cover and, thus, the same speed of the planetary gears 1704 or compound planetary gears 1704 and 1706. In some embodiments, a planetary carrier may be mechanically coupled to the main shaft 1702. In such an embodiment, a rotation of the main shaft would rotate at the same speed as the planetary carrier 1712 and, thus, the same speed of the planetary gears 1704 or compound planetary gears 1704 and 1706.
In some embodiments, a main shaft assembly 1700 may include a pump drive gear 1716. A pump drive gear may be disposed between a planetary carrier 1712 and carrier cover 1716. Further, in some embodiments, a pump drive gear 1716 may be disposed between the multiple planetary gears comprising a compound planetary gear 1704, 1706. A pump gear drive 1716 may be mechanically coupled to various components present within the gearbox assembly, including the planetary carrier 1712, planetary gears 1704, 1706, or the carrier cover 1714. A pump drive gear 1716 may interface with other components, not pictured here, within the electric engine assembly to circulate oil or other coolant liquids throughout liquid flow paths, as described herein, in an effort to cool or lubricate components present within an electric engine assembly. For example, a pump drive gear 1716 may interface with a pump gear that acts to draw liquid from a sump to a heat exchanger. In such an embodiment, the speed of rotation of the pump gear drive 1716 may determine the speed at which oil or other liquid is circulated throughout the electric engine assembly. In some embodiments, the pump drive gear 1716 may be mechanically coupled to the main shaft 1702 such that the pump drive gear rotates at the speed of the main shaft 1702. In some embodiments, main shaft assembly 1700 may comprise dowel pins 1726, or similar alignment components, that serve to align the pump drive gear with various components of the main shaft assembly 1700, including the planetary carrier 1712 or carrier cover 1714.
As described herein, an electric motor assembly may drive the rotation of a rotor. The rotation of a rotor, which may be mechanically coupled to a sun gear, may rotate the sun gear at rotor speed. The sun gear rotating at rotor speed may interface with planetary gears 1704 or compound planetary gears 1704 and 1706 to generate an output of the gearbox assembly comprising a new value of torque to be supplied to a propeller assembly. In some embodiments, the combination of using a sun gear, planetary gears, including compound planetary gears, and a ring gear, as described herein, may produce a gear reduction. Those of ordinary skill in the art would understand that a gear ratio can be calculated from the gears present in the gearbox assembly. As such, properties of the gears within the gearbox assembly may determine the gear reduction available in the electric propulsion system. In some embodiments, a gear reduction value may be a relevant design criteria for VTOL aircrafts as an aircraft may require a specific value to torque to be applied to the propeller assembly to accomplish providing lift for payloads. However, it should be understood that increasing gear size to create a larger gear reduction would result in an increase in electric engine drag profile and mass. As such, embodiments as described herein may provide optimized electric propulsion system design in terms of drag profile and mass versus payload capabilities.
As described herein, embodiments of a gearbox assembly may include a sun gear.
Embodiments of a gearbox may include a ring gear.
As described herein, embodiments of a gearbox assembly may include a planetary carrier assembly.
In some embodiments, an electric engine may include an inverter assembly. An inverter assembly may include circuitry configured to receive input of a direct current, convert the direct current to an alternating current, and provide the alternating current to the stator ring of an electric motor.
As disclosed herein, embodiments of an electric engine assembly may include a thermal management system or cooling system that may circulate a coolant or lubricant throughout the engine. A lubricant or coolant, such as oil, may reside in a sump and may be distributed to components throughout the electric engine assembly. As disclosed herein, oil may travel from a sump to a heat exchanger, to various locations in the electric engine assembly, including an inverter assembly, a gearbox assembly, and an electric motor assembly. As described herein, an electric motor assembly may include an end bell assembly. In some embodiments, an end bell assembly may abut an inverter assembly.
In some embodiments, a heat exchanger may cool oil or other liquid used to lubricate or cool the inverter assembly, gearbox assembly, and/or electric motor assembly. In some embodiments, a certain portion of the cooled oil or liquid leaving the heat exchanger may be directed to the inverter assembly to cool such components or may be directed to a motor-gearbox housing to cool components of the gearbox assembly and/or electric motor assembly. Some embodiments may include different divisions of cooled oil or liquid among the inverter assembly versus the gearbox assembly and electric motor assembly. For example, an inverter assembly may receive 40% of the cooled oil by volume and the motor-gearbox hosing may receive 60%. The ratio may differ depending on the design considerations and requirements of the particular implementation. Indeed, different types of electric propulsion systems as described herein may use different fluid distribution percentages. Further, it should also be understood that tilter electric propulsion systems and lifter electric propulsion systems may possess similar or non-similar distributions of oil from the heat exchanger. The pump corresponding to pump rotor 2104A and pump gear 2114B may provide performance improvements to a gearbox assembly. Furthermore, using the pump to drive the transportation of oil to not only a gearbox assembly and electric motor assembly, but also an inverter assembly, may eliminate the need of extra components to transport coolants to the inverter assembly. Such an advantage may reduce the mass and improving the drag profile of an electric propulsion system.
From the heat exchanger, cooled oil may enter channel 2108A, and travel, in a direction 2114A, to annulus 2110A. Annulus 2110 may be aligned along a shaft, as described herein. Annulus 2110A may include ports 2116A. Oil from channels 2108A may travel through ports 2116A to various components of the electric engine, including to a gearbox assembly and motor assembly, to provide cooling and lubrication. Oil may also travel from annulus 2110A to channel 2112A. End bell plate 2100A may also include ports 2122A that allow oil or other liquids to be transferred to through the end bell assembly 2100B. In some embodiments, the pump may create pressure, which may drive the movement of liquids through the end bell plate 2100A. For example, pressure from the pump, which may be a gerotor or positive displacement pump, may propel the travel of oil in channels 2108A, 2112A, annulus 2110A, ports 2116A, or other grooves or cavities in the end bell plate that may assist in transport of liquid.
In some embodiments, end bell assembly 2100B may comprise an end bell plate 2102B that serves to seal off an electric motor assembly housing or a motor-gearbox assembly housing. In some embodiments, an end bell assembly 2100B may comprise a first circular wall extending away from the end bell plate 2102B. In some embodiments, a ring gear 2106B may be coupled to the first circular wall 2104B such that the ring gear 2106B is not free to rotate, as described herein. In some embodiments, an end bell assembly 2100B may comprise a second circular wall 2108B extending away from the end bell plate 2102B. In some embodiments, the second circular wall 2108B may possess a diameter that is less than a diameter of the first circular wall 2104B. A second circular wall 2108B may housing a bearing 2110B. In some embodiments, the bearing 2110B may be mechanically coupled to a shaft, including a main shaft that may transfer mechanical shaft power to a propeller assembly. In some embodiments, bearing 2110B may include grooves to assist in the transfer of oil or other liquids. The second circular wall 2108B may also comprise an annulus that includes port holes 2112B. Port holes 2112B may aligned with ports 2116A to receive oil or liquid from the heat exchanger. Port holes 2112B may comprise a supply of oil or other liquid to cool or lubricate components of the electric motor assembly and gearbox assembly.
In some embodiments, the port holes 2112B may transfer oil or other liquid to the main shaft. In some embodiments, an outer surface of a main shaft may serve as a liquid flow path where the oil or other liquid flows upon the main shaft and may be distributed to components within the gearbox assembly and/or electric motor assembly.
Disclosed embodiments of an inverter assembly may include an inverter assembly with a heat exchanger.
In some embodiments, an inverter assembly 2200 may comprise a high voltage connector 2212 and low voltage connectors 2210. High voltage connector 2212 may have a low profile. High voltage connector 2212 may receive high voltage power from a high voltage power system located elsewhere within the aircraft via high voltage channels. Inverter assembly 2212 may include at least one drain 2208. Drains 2208 may be configured to allow any oil or liquid present within the inverter assembly to exit the inverter assembly 2200 no matter the orientation of the electric engine assembly. In alternative embodiments, inverter assembly 2200 may also include vents. In some embodiments, an inverter assembly 2200 may comprise a heat exchanger 2206 coupled or mounted to the thermal plate 2204. In some embodiments, heat exchanger 2206 may be an integrated heat exchanger. In some embodiments, thermal plate 2204 may be welded to heat exchanger 2206. For an example, thermal plate 2204 may be comprised of aluminum. Assembly of thermal plate 2204 and heat exchanger 2206 may include brazing, quenching, aging, and welding. In some embodiments, thermal plate 2204 and heat exchanger 2206 may be machined from the same material.
Some embodiments may include an inverter assembly wherein the components of the inverter abut one another and may share a common housing. In some embodiments, components of the inverter assembly can be placed on top of one another in a stacked orientation. In some embodiments, components of the inverter assembly can be substantially aligned along a central axis. An inverter assembly may include various components for sensing, circuitry, and controls.
In some embodiments, the stacked orientation in the housing may conform to various design shapes, for example a circular shape possessing a diameter proportional to that of a motor or gearbox, or any other design shapes. The internal components of the inverter may be arranged to assist in achieving that design goal shape. In some embodiments, a stacked orientation may be achieved by using common structural components through the stack, for example designing the different levels of the stack such that a common structure, such as various bolts of the same length, can pass through each level to create a stacked orientation.
In some embodiments, using a stacked orientation may create additional obstacles with respect to additional design considerations, such as heat transfer where the difficulty of managing the proper distribution of coolant could increase in such a configuration. Further, using long bolts to travel through various levels of an inverter my increase the shock and vibrations experienced by the inverter assembly. However, it is also understood that such a stacked orientation may be advantageous in view of various design considerations. For example, allowing a stacked orientation may be beneficial from an aerodynamic perspective where the stacking allows the inverter, or the inverter in combination with other engine components such as the gearbox and/or motor, to maintain a low drag profile. Additionally, a stacked orientation may be advantageous from a manufacturability point of view such that less components are involved in securing the inverter assembly, as well as being advantageous from a mass reduction perspective where less components, and potentially less mass, is being used to secure the components.
As disclosed herein, an inverter assembly may include a power PCBA assembly. In some embodiments, power PCBA assembly may include a power board.
As discussed above, an electric engine and related control components of a VTOL aircraft may generate heat during operation. For example, such components may include an inverter assembly, electric motor assembly, and a gearbox assembly. An engine may accumulate a buildup of heat generated from mechanical friction between parts, and from resistive heating within the motor-gearbox assembly. The accumulated heat may be carried to the heat exchanger by lubricant circulating through one or more parts of the engine. The heat must be dissipated to prevent degradation or damage to the motor, control components and other elements of the VTOL aircraft. Such heat may be managed by cooling the engine, including by direct or indirect cooling. In some embodiments, cooling may be assisted by a heat exchanger. A heat exchanger may be configured to receive a circulating heat exchange medium from an electric engine. For example, the heat exchange medium may comprise oil, and the oil may be used to both lubricate and cool the components of the electric engine. A heat exchanger may interface one or more fluids with each other, thereby cooling a fluid that is at a higher temperature. A heat exchanger may be advantageously located next to an electric engine, thereby minimizing the volume (and weight) of material required to achieve the cooling and lubricating functions. In some embodiments, a heat exchanger may be fluidically, thermally, and mechanically coupled to an inverter assembly such that the heat exchanger may share common connections with the inverter assembly, reducing the need for components such as cables, wires, tubes, and hoses, which may add weight and require more space in the electric engine. Heat in the inverter assembly, electric motor assembly, or gearbox assembly may be transferred to a cooling fluid such as oil. The oil may absorb such heat, and the oil may then be directed to a sump.
As described herein, an electric propulsion system may include a heat exchanger.
As described herein, embodiments of an electric engine may include circulating a lubricant or coolant throughout the engine. A heat exchanger may cool lubricants or coolants directed to engine components such as a motor, gearbox, or inverter. In some embodiments, heat from the inverter assembly may be directly conducted to the coolant or lubricant, such as oil.
In some embodiments, a lubricant or coolant may be used to cool an inverter assembly using a thermal plate 2702B. Some embodiments may include hot oil or other liquid entering inlet channel 2708A, that may be aligned with heat exchanger inlet 2706B, to enter the heat exchanger 2710B. Cooled oil or other liquid exiting the outlet channel 2710A aligned with heat exchanger outlet 2708B and a portion, or all, of the cooled oil or liquid may follow a liquid flow path to an annulus 2712B of the thermal plate. Oil or other liquid in annulus 2712B may be distributed to heat sinks located on the thermal plate 2702B such as fin arrays 2714B and may provide cooling to power modules, such as those referenced in
Disclosed embodiments of the electric propulsion system may include one or more components for distributing a lubricant or coolant, as described herein. In some embodiments, the lubricant or coolant, such as oil, may be circulated from a sump 1212A, to a heat exchanger 1226A, and then to an inverter assembly, gearbox assembly, and motor assembly (as illustrated in
As described herein, by using a fluid, such as oil, to provide both lubrication and cooling, the amount of oil used in the electric propulsion system may be minimized. Furthermore, the oil can be used for lubricating various bearings, such as rolling bearings or hydrodynamic bearings, as described herein. Minimizing the amount of oil used in the electric propulsion system may reduce the mass and drag profile of the electric propulsion system. Furthermore, minimizing the amount of oil necessary for operation of the electric propulsion system may enable the total quantity of oil in the electric propulsion system to remain below a threshold amount. For example, the electric propulsion system may reduce the amount of oil necessary for operation by using a heat exchanger, as described herein. Warm or hot oil that has been used for lubrication and cooling may reside in a sump, and by using a heat exchanger to cool such oil, the electric propulsion system may re-use the oil, eliminating the need for additional oil. Further, as described herein, the use of a common liquid for both cooling and lubricating can lead to a reduction in mass of components when compared to other methods of cooling and lubricating using various liquids. Such a configuration using various liquids may require additional mass and size of the electric propulsion system such as additional heat exchangers, additional fluid distribution channels or tubes, and additional surface area to receive cooling air from a propeller assembly.
In some embodiments, ventilation from air-flow may provide cooling to lubricants within the heat exchanger. It is understood that by using oil to not only lubricate the electric engine but also cool the electric engine rather than another coolant, additional oil will be added to the system, but that oil will remove traditional components that may be used to cool such an electric engine. For example, if the electric engine were cooled by another liquid such as glycol, the engine may comprise separate heat exchangers for both the lubricant fluid and the coolant fluid. As such, in embodiments where a single fluid is being used for both lubrication and cooling, such as oil, an increase in oil would be present but there would only be a need for one heat exchanger, so there may be a decrease in mass, due to using less heat exchangers and potentially other components not being required, of the overall system and a more appealing drag profile may be present. Further, using one substance for the lubrication and cooling of the engine may increase efficiency of the system due to the reduction in mass and the benefits of cooling the engine with a substance rather than relying on air cooling which may have issues traveling throughout the engine.
Some embodiments of an inverter may include an inverter possessing a coolant path traveling around the outer edge of the inverter but within the inverter housing rather than utilizing a heat exchanger. For example, a coolant path may travel around any printed circuit board assemblies, power modules, or other inverter components present in the inverter.
As disclosed herein, embodiments of electric engine may include an inverter assembly. In some embodiments, the inverter assembly may include a thermal plate.
As described above, a tilter may possess a variable pitch mechanism that serves to change the pitch of the propeller blades of a VTOL aircraft. In some embodiments, a variable pitch mechanism may be mounted to the rear of the electric engine assembly, such as the rear of an inverter assembly. Further, a variable pitch mechanism may interact with a main shaft, as described herein, to alter the pitch of the propeller blades. In such embodiments, the inverter assembly 3104A, inverter assembly housing 3116A, and divider plate, as discussed herein, may possess a packaging having a passage through their configurations and housings to allow the variable pitch mechanism to interface with the main shaft or propeller blades. As discussed throughout, a lifter electric propulsion system may not alter its orientation of thrust or pitch of blades. Therefore, in some embodiments, a divider plate may not possess a passage such as the one present in the tilter electric propulsion system. Further, the inverter assembly and inverter assembly housing of a lifter electric propulsion system may not possess such a passage, but it is recognized that from a safety testing point of view and a manufacturability standpoint, it may be beneficial to have the inverter assembly and inverter assembly housing of a lifter electric propulsion system possess a similar packaging, including the passage, to that of the tilter electric propulsion system.
As discussed herein, it is noted that having similar components between the tilter and lifter electric propulsion systems may be beneficial with respect to manufacturability of the overall aircraft. Further, using similar components between the tilter and lifter electric propulsion systems may be beneficial in terms of diagnosing issues and assuring safety requirements and protocols are met. However, in some embodiments, a lifter and tilter may possess components that are not present within the other. For example, a lifter electric propulsion system 3100B may include a lock nut 3112B posited between the main shaft 3108B and the shaft flange assembly 3120B that is larger than the lock nut present within the tilter electric propulsion system 3100A. A lock nut 3122B may serve to ensure the mechanical coupling of the main shaft 3108B and shaft flange assembly 3120B may not be damaged or corrupted due to the various vibrations loads experienced throughout the flight. For example, as discussed herein, some phases of flight do not require the lifter electric propulsion system to be active and in such cases may require the blades to be stored in a certain fashion. However, if the lifter blades were to not be properly stored, they may experience a drag force against the blades and the mechanical coupling of the main shaft 3108B and shaft flange assembly 3122B may experience a tension force. Further, in some embodiments, the lock nut 3112B of the lifter electric propulsion system 3100B may counteract operational loads. In some embodiments, a lifter electric propulsion system may also include a larger propeller flange 3126A, compared to the shaft flange of the tilter, for similar reasons as to the presence of the lock nut 3122B. Further, the lifter electric propulsion system may also include a bearing 3124A to assist in the rotation of the propeller flange 3126A. As described herein, an electric propulsion system may achieve different angles of orientation during operation. As such, fluids in the electric propulsion system, including coolants or lubricants, may move due to gravitational forces. For example, a lubricant or coolant such as oil may be shifted within the electric propulsion system during operation. Oil may reside in a sump, and the oil may shift within the sump and the electric propulsion system. Not matter the orientation, some embodiments may require some quantity of oil or other liquid acting as coolant or lubricant throughout all phases of flight. As such, a cooling system may be designed to allow for the circulation of oil no matter the orientation of the aircraft.
It is understood that by using oil to not only lubricate the electric engine but also cool the electric engine rather than another coolant, additional oil will be added to the system, but that oil will remove traditional components that may be used to cool such an electric engine. For example, if the electric engine were cooled by another liquid such as glycol, the engine may comprise separate heat exchangers for both the lubricant fluid and the coolant fluid. In some embodiments, an electric engine may be cooled using various liquids. Some embodiments may include the electric propulsion system comprising multiple heat exchangers that cool their respective liquid flowing in their respective liquid flow paths. In some embodiments, multiple cooling and/or lubricating liquids, such as glycol and oil, may have a respective liquid flow path that circular through a common heat exchanger, also known as a dual heat exchanger. In such a configuration, the number of heat exchangers may be less than the number of types of liquid flow paths, based on liquid type, and the overall propulsion system may conserve mass by not possessing multiple, or more, heat exchangers.
However, with respect to embodiments utilizing a single fluid for both lubrication and cooling, such as oil, an increase in oil would be present but there would only be a need for one heat exchanger, so there may be a decrease in mass, due to using less heat exchangers and potentially other components not being required, of the overall system and a more appealing drag profile may be present. Further, using one substance for the lubrication and cooling of the engine may increase efficiency of the system due to the reduction in mass and the benefits of cooling the engine with a substance rather than relying on air cooling which may have issues traveling throughout the engine.
With respect to using oil, or other flammable liquids, within the electric propulsion systems as described herein, federal laws and regulations may be in place requiring safety components such as fire protective barriers adjacent to engines that use more than a threshold amount of oil or other flammable materials. Such federal laws and regulations may be enforced by government entities such as the Federal Aviation Administration.
It should be noted that some of the electric propulsion system embodiments as described herein may not include a fire protective barrier. A fire protective barrier as used herein may include an engine component or aircraft component designed, constructed, or installed with the primary purpose of preventing a hazardous quantity of air, fluid, or flame from passing around or through the fire protective barrier, and/or to protect against corrosion. In some embodiments, a fire protective barrier may be required for each electric propulsion system present on an aircraft. As such, if an aircraft, as described herein, possesses, for example, twelve electric propulsion systems, twelve fire protective barriers may be required to be installed on the aircraft. In some embodiments, a fire protective barrier may be required to be present on each wing, to surround the fuselage, or any other configuration based on federal law, regulations, or other safety requirements. As such, the presence of a fire protective barrier may add additional mass to the aircraft and thus decrease the efficiency of the electric propulsion system and further limit reduce the amount of payload, including passengers, present on the aircraft. This is especially relevant for VTOL aircraft design where a single aircraft may possess, for example, twelve electric propulsion systems so any increase in mass due to a single fire protective barrier could be experienced twelve-fold.
While some embodiments as described herein do not include a fire protective barrier, additional versions all embodiments described and considered herein may include a fire protective barrier. In addition, each embodiment as described herein may possess various types and locations of fire protective barriers.
In some embodiments, fire risk management in the aircraft design may not be limited to inclusion of a fire protective barrier. Additional design considerations may address fire risks, such as additional components to ensure an aircraft may maintain flight if a fire were to occur. For example, an aircraft boom, as described herein, may feature additional components present within the boom to ensure that if a fire was present, components were lost to a fire, components became detached due to fire, or any other loss in functionality or components were to occur, the aircraft may still maintain balanced flight.
As discussed above and throughout this disclose, an exemplary electric propulsion system may include components comprising an electric motor assembly, a gearbox assembly, and an inverter assembly across various configurations, such as representative configurations as described herein. Exemplary embodiments as discussed herein may include components of the electric propulsion system being aligned along a common axis or substantially aligned along a common axis. In some embodiments the components may be aligned along a shaft or main shaft that provides mechanical shaft power to turn the propellers of a propeller assembly. Some embodiments may include components of the electric propulsion system abutting each other in a sequence along an axis or substantially aligned along an axis. In some embodiments, the location or positioning of one or more electric propulsion system components may provide a reduction in mass of the system and yield a more efficiency drag profile. For example, one or more component may be substantially aligned along a common axis or abut each other, and may not require additional components such as connection wires or additional housing volume to allow for wires. As such, the respective mass and volume typically needed to allow such connection wires may not be needed in the disclosed electric propulsion systems.
In some embodiments, an electric propulsion system may also comprise a cooling system configured to target multiple heat-generating portions of electric propulsion system. Some embodiments may include portions of the electric propulsion system being air-cooled by air flow generated from the propeller assembly or air flow that is encountered during various phases of flight. Some embodiments may include portions of the electric propulsion system being cooled using one or more liquid flow paths throughout the electric propulsion system. Such embodiments may also include the liquid flow paths circulating through a heat exchanger that is exposed to air flow such that any heat contained in the liquid flow paths may be transferred to the air flowing through the heat exchanger. It should be understood that the components of the electric propulsion system, as described herein, may all be cooled using a common cooling system, may each have their own independent cooling system, or may combine various types and configurations of cooling systems. In some embodiments, the respective cooling system of an electric propulsion system may have an impact on efficiency of the components of the electric propulsion system. For example, in some embodiments, liquid cooling may allow for an inverter assembly to operate more efficiently than an inverter assembly utilizing an air-cooled system.
In some embodiments, an electric propulsion system may include a cooling system utilizing liquid cooling. In some embodiments, a cooling system liquid may include glycol, oil, or any other liquid that enables the transfer of heat from components of the electric propulsion system to the liquid. Further, some embodiments may include cooling an electric propulsion system using a liquid that is also used for lubricating components of the electric propulsion system. In some embodiments, the electric propulsion system may include a cavity, reservoir, or sump for collecting and circulating coolant liquid throughout the electric propulsion system.
As shown in
In some embodiments, an electric propulsion system may include components that are not aligned, or substantially aligned, along an axis. For example, an electric propulsion system may include an electric motor assembly aligned along a shaft that provides mechanical shaft power to the propeller assembly and an inverter assembly supplying alternating current to the electric motor assembly that is located elsewhere within the aircraft. Some embodiments may include an inverter assembly that does not abut the electric motor assembly but is instead housed elsewhere within the boom, wing, or fuselage. In such embodiments, wiring may be run from the inverter assembly to the electric motor assembly to transmit alternating current from the inverter. Separating the locations of components of the electric propulsion system may lead to an increase in mass of the aircraft due to the required wiring and other connection components.
In some embodiments, an electric propulsion system may include thermal management, also referred to as cooling systems herein, that include liquid cooling. As disclosed herein, some exemplary cooling systems may include distributing a liquid coolant to components located throughout the electric motor assembly, the gearbox assembly, and the inverter assembly. However, it should be understood that cooling systems as disclosed herein may also include liquid coolant that is circulated about the perimeter of the electric motor assembly, the gearbox assembly, and/or the inverter assembly. For example, a cooling system may comprise a cavity, jacket, or distribution channels of a cooling system that circulates liquid coolant about the perimeter of components located within the electric propulsion system.
As discussed herein, an electric propulsion system may include various configurations of components such as aligned along an axis, abutting one another, substantially aligned along an axis, or components connected using wires or other methods of connection. As such, some embodiments may include components sharing a housing. For example, as discussed above, a gearbox assembly may be housed within a motor assembly housing. Further, some embodiments may include housing a gearbox assembly, inverter assembly, or other assemblies or components of those assemblies within the propeller assembly. Such configurations may be driven by design constraints such as weight, drag profile, lift, torque, payload, flight time, or any other design constraints relevant to VTOL aircrafts.
In some embodiments, an electric propulsion system may include components present within various component housings. As discussed herein, various components of the electric propulsion system may be present within housings and may be organized in various ways within those housings. Some embodiments of an electric propulsion system may include various configurations of components to achieve varying design goals. Differing embodiments may possess differing primary design components that must be achieved at the expense of other design criteria. For example, some embodiments may contain redundant systems that may add extra mass to the aircraft yet increase passenger safety by avoiding and/or removing single points of failure. Further, some embodiments of an electric propulsion system may include various types of thermal management systems, also referred to herein as cooling systems. Some electric propulsion systems may include a combination of cooling systems, such as air-cooling and liquid cooling systems. For example, electric propulsion system components may possess air-cooling designs such as components being mechanically coupled to cooling fins and liquid cooling designs where components are fluidically coupled to liquid flow paths and a heat exchanger that extracts heat from the liquid and transfers it into external air.
The liquid paths 4016C and 4016D are illustrated with a high level of generality as a simple loop. However, it should be understood that the liquid paths may comprise branches, sub-loops or other segmented paths. In general, the liquid may be circulated in any way that effectively lubricate and cool various components present within the motor-gearbox housing 4012C and 4012D.
The liquid flow paths 4418A, 4418B, and 4418C are illustrated with a high level of generality as a simple loop. In some embodiments, liquid flow paths may comprise branches, sub-loops or other segmented paths. In general, the liquid may be circulated in any way that effectively lubricate and cool various components present within the motor-gearbox housings 4414A-C and inverter assembly housings 4416A-C.
The liquid paths 4616A and 4616B are illustrated with a high level of generality as a simple loop. In some embodiments, the liquid paths may comprise branches, sub-loops or other segmented paths. In general, the liquid may be circulated in any way that effectively lubricate and cool various components present within the motor assembly housing 4612A and 4612B.
The embodiments may further be described using the following clauses:
Clause Set A: 1. An electric engine for a vertical takeoff-and-landing aircraft, comprising: an electric motor assembly including a stator and a rotor; an inverter assembly; a gearbox assembly including: a sun gear; a main shaft including a length of the main shaft that extends from a first end of the main shaft through the gearbox assembly and through the electric motor assembly to a second end of the main shaft; and a hydrodynamic bearing located between the main shaft and sun gear.
2. The electric engine of clause A1, further comprising a bearing including an inner race mechanically coupled to the main shaft and an outer race mechanically coupled to the rotor.
3. The electric engine of clause A2, wherein the bearing comprises a deep groove ball bearing.
4. The electric engine of clause A2, wherein the bearing comprises a rolling bearing.
5. The electric engine of clause A1, wherein the hydrodynamic bearing includes oil.
6. The electric engine of clause A1, further comprising a heat exchanger including one or more liquid flow paths fluidically coupled to the gearbox assembly and electric motor assembly.
7. The electric engine of clause A6, wherein the liquid flow paths are configured to allow flow of oil.
8. The electric engine of clause A6, wherein the hydrodynamic bearing comprises liquid from the one or more liquid flow paths.
9. The electric engine of clause A1, wherein the main shaft travels through a hollow sun gear.
10. The electric engine of clause A9, wherein the hydrodynamic bearing is disposed between outer surface of the main shaft and an inner surface of the hollow sun gear.
11. The electric engine of clause A2, wherein the bearing is mechanically coupled to a flange configured to rotate with the shaft, wherein the flange is mechanically coupled to the main shaft.
12. The electric engine of clause A2, wherein the bearing is configured to support the rotor.
13. The electric engine of clause A1, wherein the hydrodynamic bearing is configured to counteract a rotor moment load experienced by the sun gear.
14. The electric engine of clause A1, wherein the gearbox assembly further comprises a planetary gear.
15. The electric engine of clause A1, wherein the gearbox assembly further comprises a ring gear.
16. The electric engine of clause A1, wherein the gearbox assembly further comprises a planetary carrier.
17. An electric engine for a vertical takeoff-and-landing aircraft, comprising: an electric motor assembly including a stator and a rotor; an inverter assembly; a gearbox assembly including: a sun gear; and a bearing including an outer race mechanically coupled to an inner surface of the rotor.
18. The electric engine of clause A17, wherein the sun gear is disposed concentric with the rotor.
19. The electric engine of clause A18, wherein the bearing is configured to guide a rotation of the sun gear and the rotor.
20. The electric engine of clause A18, wherein the outer race of the bearing is mechanically coupled to an inner surface of a hollow sun gear.
21. The electric engine of clause A20, further comprising a shoulder of a first end of the sun gear disposed concentric with a shoulder of a first end of the rotor, wherein the shoulders cooperate to restrict movement of the bearing in an axial direction of the shaft.
22. The electric engine of clause A20, wherein a diameter of the outer race of the bearing is substantially the same as a diameter of the inner surface of a first end of the sun gear and a diameter of the inner surface of a first end of the rotor.
23. The electric engine of clause A20, wherein a first end of the sun gear and a first end of the rotor meet on the outer race of the bearing.
24. The electric engine of clause A17, wherein the outer race of the bearing is configured to align a first end of the sun gear with a first end of the rotor such that the first end of the sun gear abuts the first end of the rotor.
25. The electric engine of clause A17, wherein the bearing comprises a rolling bearing.
26. The electric engine of clause A18, wherein the sun gear and rotor are removably attached.
27. The electric engine of clause A17, wherein the gearbox assembly further comprises a planetary gear.
28. The electric engine of clause A17, wherein the gearbox assembly further comprises a ring gear.
29. The electric engine of clause A17, wherein the gearbox assembly further comprises a planetary carrier.
30. An electric engine for a vertical takeoff-and-landing aircraft, comprising: an inverter assembly; a gearbox assembly including: a sun gear; an electric motor assembly; a main shaft including a length of the main shaft that extends from a first end of the main shaft through the gearbox assembly and through the electric motor assembly to a second end of the main shaft; and a hydrodynamic bearing located between the main shaft and sun gear, wherein the inverter assembly, the gearbox assembly, and the electric motor assembly are substantially aligned along an axis and each abuts at least one of the others.
31. An electric engine for a vertical takeoff-and-landing (VTOL) aircraft comprising the aircraft of any clauses A1-A30.
Clause Set B: 1. A method for balancing a rotor of an electrical engine of an electrical propulsion system comprising: identifying an axis of rotation of a rotor, wherein the rotor comprises a sacrificial layer having a mass M formed along a circumference of the rotor; determining an imbalance present in the rotor by rotating the rotor about the axis of rotation, wherein determining an imbalance includes marking the rotor, rotating the rotor, and detecting an amplitude of the imbalance; calculating an amount of mass k to add or remove at a position p along the sacrificial layer such that a center of mass of the rotor coincides with the axis of rotation of the rotor; and removing an amount of mass r from the sacrificial layer such that an amount of remainder mass n is present along the circumference of the rotor.
2. The method of clause B1, wherein the amount of remainder mass n is equal to the amount of mass k.
3. The method of clause B1, further comprising a thickness of the sacrificial layer based on the precision of the machine machining the rotor.
4. The method of clause B1, wherein the rotor possesses a number of sacrificial layers greater than two.
5. The method of clause B1, wherein rotating the rotor about the axis of rotation includes rotating at a speed less than operating speed.
6. The method of clause B1, wherein rotating the rotor about the axis of rotation includes rotating at a speed less than the first resonance of the rotor.
7. The method of clause B1, wherein the remainder mass n is present at the position p when the amount of mass k is calculated to be added.
8. The method of claim 1, wherein the remainder mass n is present at a position q when the amount of mass k is calculated to be removed, wherein the position q is on the opposite side of the sacrificial layer from the position p.
9. The method of clause B1, wherein removing the amount of mass r includes an amount of mass equal to or less than M-k.
10. The method of clause B1, wherein detecting the amplitude of the imbalance may include using a machine to track the rotor mark during rotation and calculate the amplitude of the imbalance.
11. The method of clause B10, wherein the machine to track the rotor mark during rotation and calculate the amplitude of the imbalance may include an electric eye.
12. The method of clause B10, wherein the machine to track the rotor mark during rotation and calculate the amplitude of the imbalance may include an encoder.
13. The method of clause B1, wherein removing includes machining away.
14. The method of clause B1, wherein removing include using a machine with a milling resolution that is less than or equal to five microns.
15. The method of clause B1, wherein the rotor is made of aluminum.
16. The method of clause B1, wherein the sacrificial layer is made of aluminum.
17. The method of clause B1, further comprising a first end of the rotor and a second end of the rotor, wherein each end comprises a substantially circular profile, wherein a first sacrificial layer is posited a first distance from the first end of the rotor and a second sacrificial layer is posited a second distance from the second end of the rotor.
18. The method of clause B17, wherein the first sacrificial layer and the second sacrificial layer have substantially similar depths and widths.
19. The method of clause B17, wherein the first distance is substantially similar to the second distance.
20. The method of clause B17, wherein position p is located along either the first sacrificial layer or second sacrificial layer.
21. The method of clause B1, wherein the sacrificial layer includes N grooves defining N sacrificial portions, wherein the grooves are configured to guide oil flow through the rotor in normal operation.
22. The method of clause B1, wherein removing an amount of mass r from the sacrificial layer includes removing at least 50% of the volume of the sacrificial layer.
23. The method of clause B1, wherein removing an amount of mass r from the sacrificial layer includes removing at least 60% of the mass of the sacrificial layer.
24. The method of clause B1, wherein removing an amount of mass r from the sacrificial layer includes removing at least 70% of the mass of the sacrificial layer.
25. The method of clause B1, wherein removing an amount of mass r from the sacrificial layer includes removing at least 75% of the mass of the sacrificial layer.
26. The method of clause B1, wherein removing an amount of mass r from the sacrificial layer includes removing at least 80% of the mass of the sacrificial layer.
27. The method of clause B1, wherein removing an amount of mass r from the sacrificial layer includes removing at least 90% of the mass of the sacrificial layer.
28. The method of clause B1, wherein removing an amount of mass r from the sacrificial layer includes removing at least 100% of the mass of the sacrificial layer.
29. The method of claim 1, wherein removing an amount of mass r from the sacrificial layer includes using a machine capable of removing less than 0.1% of the sacrificial layer material by volume.
30. A method for balancing a rotor of an electrical engine of an electrical propulsion system comprising the method of any clauses B1-B29.
Clause Set C: 1. A method for balancing a rotor assembly of an electrical engine of an electrical propulsion system comprising: identifying an axis of rotation of a rotor, wherein the rotor comprises a sacrificial layer having a mass M formed along a circumference of the rotor; determining an imbalance present in the rotor by rotating the rotor about the axis of rotation, wherein determining an imbalance includes marking the rotor, rotating the rotor, and detecting an amplitude of the imbalance; calculating an amount of mass k to add at a position p along the sacrificial layer such that a center of mass of the rotor coincides with the axis of rotation of the rotor; removing an amount of mass r from the sacrificial layer such that an amount of remainder mass n is present along the circumference of the rotor; identifying an axis of rotation of a rotor assembly, wherein the rotor assembly comprises the rotor mechanically coupled to a sun gear; determining an imbalance present in the rotor assembly by rotating the rotor assembly about the axis of rotation, wherein determining an imbalance includes marking the rotor assembly, rotating the rotor assembly, and detecting an amplitude of the imbalance; calculating a number j of rivets to add to the rotor assembly such that the center of mass of the rotor assembly coincides with the axis of rotation of the rotor assembly; and adding the j rivets to the rotor assembly.
2. The method of clause C1, wherein the amount of remainder mass n is equal to the amount of mass k.
3. The method of clause C1, further comprising a thickness of the sacrificial layer based on the precision of the machine machining the rotor.
4. The method of clause C1, wherein the rotor possesses an even number of sacrificial layers greater than two.
5. The method of clause C1, wherein rotating the rotor about the axis of rotation includes rotating at a speed less than operating speed.
6. The method of clause C1, wherein rotating the rotor about the axis of rotation includes rotating at a speed less than the first resonance of the rotor.
7. The method of clause C1, wherein the remainder mass n is present at the position p.
8. The method of clause C1, wherein the imbalance occurs when the axis of rotation does not align with a center of mass of the rotor.
9. The method of clause C1, wherein removing the amount of mass r includes an amount of mass equal to or less than M-k.
10. The method of clause C1, wherein detecting the amplitude of the imbalance may include using a machine to track the rotor mark during rotation and calculate the amplitude of the imbalance.
11. The method of clause C10, wherein the machine to track the rotor mark during rotation and calculate the amplitude of the imbalance may include an electric eye.
12. The method of clause C10, wherein the machine to track the rotor mark during rotation and calculate the amplitude of the imbalance may include an encoder.
13. The method of clause C1, wherein removing includes machining away.
14. The method of clause C1, wherein the rotor is made of aluminum.
15. The method of clause C1, wherein the sacrificial layer is made of aluminum.
16. The method of clause C1, further comprising a first end of the rotor and a second end of the rotor, wherein each end comprises a substantially circular profile, wherein a first sacrificial layer is posited a first distance from the first end of the rotor and a second sacrificial layer is posited a second distance from the second end of the rotor.
17. The method of clause C1, wherein the first sacrificial layer and the second sacrificial layer have substantially similar depths and widths.
18. The method of clause C1, wherein the first distance is substantially similar to the second distance.
19. The method of clause C1, wherein position p is located along one of the first sacrificial layer or second sacrificial layer.
20. The method of clause C1, wherein the sacrificial layer includes N grooves defining N sacrificial portions, wherein the grooves are configured to guide oil flow through the rotor in normal operation.
21. The method of clause C1, wherein rotating the rotor assembly about the axis of rotation includes rotating at a speed less than operating speed.
22. The method of clause C1, wherein rotating the rotor assembly about the axis of rotation includes rotating at a speed less than the first resonance of the rotor assembly.
23. The method of clause C1, wherein detecting the amplitude of the imbalance may include using a machine to track the rotor assembly mark during rotation and calculate the amplitude of the imbalance.
24. The method of clause C23, wherein the machine to track the rotor assembly mark during rotation and calculate the amplitude of the imbalance may include an electric eye.
25. The method of clause C23, wherein the machine to track the rotor assembly mark during rotation and calculate the amplitude of the imbalance may include an encoder.
26. The method of clause C1, wherein the sun gear is disposed concentric with the rotor.
27. The method of clause C1, wherein the sun gear and rotor are removably attached.
28. The method of clause C1, wherein the rotor possesses a circular layer comprising a plurality of through holes.
29. The method of clause C28, wherein the rivets are affixed to the plurality of through holes.
30. The method of clause C28, wherein the sun gear and rotor are removably attached via the plurality of through holes.
31. The method of clause C1, wherein the rivets are aluminum.
32. The method of clause C1, wherein the rivets are steel.
33. The method of clause C1, wherein the rivets are copper.
34. The method of clause C1, wherein at least two rivets are of different materials.
35. The method of clause C1, wherein the number j of rivets is an integer.
36. The method of clause C1, wherein the rotor assembly includes the rotor having the remainder mass n present along the circumference of the rotor.
37. The method of clause C1, wherein removing an amount of mass r from the sacrificial layer includes removing up to 50% of the mass of the sacrificial layer.
38. The method of clause C1, wherein removing an amount of mass r from the sacrificial layer includes removing up to 60% of the mass of the sacrificial layer.
39. The method of clause C1, wherein removing an amount of mass r from the sacrificial layer includes removing up to 70% of the mass of the sacrificial layer.
40. The method of clause C1, wherein removing an amount of mass r from the sacrificial layer includes removing up to 75% of the mass of the sacrificial layer.
41. The method of clause C1, wherein removing an amount of mass r from the sacrificial layer includes removing up to 80% of the mass of the sacrificial layer.
42. The method of clause C1, wherein removing an amount of mass r from the sacrificial layer includes removing up to 90% of the mass of the sacrificial layer.
43. The method of clause C1, wherein removing an amount of mass r from the sacrificial layer includes removing up to 100% of the mass of the sacrificial layer.
44. The method of clause C1, wherein the remainder mass n is present at a position q when the amount of mass k is calculated to be removed, wherein the position q is on the opposite side of the sacrificial layer from the position p.
45. The method of clause C1, wherein removing include using a machine with a milling resolution that is less than or equal to five microns.
46. A method for balancing a rotor assembly of an electrical engine of an electrical propulsion system comprising the method of any clauses C1-C45.
Clause Set D: 1. An electrical propulsion system having a gearbox apparatus that delivers power from an electrical engine via a reverse torque path, the gearbox apparatus comprising: a planetary gear mechanically coupled to a rotor of an electric motor; a planetary carrier that is connected to at least one shaft from a set of shafts that extends concentrically from the at least one planetary gear; a main shaft comprising: a first end of the main shaft extending through the planetary carrier, a second end of the main shaft mechanically coupled to a propeller assembly, and a length of the main shaft extending the first end of the main shaft through a sun gear and rotor to a second end of the main shaft; and a carrier cover that is connected to the main shaft and is connected to at least one shaft from a set of shafts that extends concentrically from the at least one planetary gear.
2. The gearbox apparatus of clause D1, wherein a rotation of the carrier cover drives a rotation of the main shaft.
3. The gearbox apparatus of clause D1, wherein the sun gear mechanically coupled to the rotor of the electric motor, wherein the electrical motor further comprises a stator.
4. The gearbox apparatus of clause D1, wherein the sun gear is hollow.
5. The gearbox apparatus of clause D1, further comprising a ring gear that interfaces with the at least one planetary gear.
6. The gearbox apparatus of clause D1, wherein the planetary gear comprises a compound planetary gear.
7. The electrical propulsion system of clause D1, wherein the gearbox comprises a multi-stage planetary gearset.
8. The electrical propulsion system of clause D1, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 1 quart.
9. The electrical propulsion system of clause D1, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 1.5 quarts.
10. The electrical propulsion system of clause D1, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 2 quarts.
11. The electrical propulsion system of clause D1, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 2.5 quarts.
12. The electrical propulsion system of clause D1, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 3 quarts.
13. The electrical propulsion system of clause D1, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 5 quarts.
14. The gearbox apparatus of clause D1, further comprising a gear ratio that is about 6.45.
15. The gearbox apparatus of clause D1, wherein the sun gear, the planetary gear, the ring gear, the planetary carrier, and the carrier cover are disposed concentric with the main shaft.
16. A method of delivering power from an electrical engine using a gearbox via a reverse torque path, comprising: driving a planetary gear that is mechanically coupled to a rotor of an electric motor, wherein the at least one planetary gear is driven by the sun gear and interfaces with a ring gear; driving a planetary carrier that is connected to at least one shaft that extends concentrically from the planetary gear; driving a carrier cover that is connected to at least one shaft from a set of shafts that extends concentrically from the planetary gear; and driving a main shaft comprising driving a first portion of the main shaft, wherein the first portion of the main shaft is mechanically coupled to the carrier cover, and transferring torque along the main shaft to a second portion of the main shaft, wherein the second portion of the main shaft is mechanically coupled to a propeller assembly.
17. The method of clause D16, further comprising driving a sun gear that is mechanically coupled to the rotor of the electric motor, wherein the electrical motor further comprises a stator.
18. The method of clause D16, wherein the sun gear is hollow.
19. The method of clause D16, wherein the planetary gear includes a compound planetary gear.
20. The method of clause D16, further comprising a ring gear that interfaces with the planetary gear.
21. The method of clause D20, wherein the ring gear is fixed.
22. The method of clause D16, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 1 quart.
23. The method of clause D16, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 1.5 quarts.
24. The method of clause D16, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 2 quarts.
25. The method of clause D16, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 2.5 quarts.
26. The method of clause D16, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 3 quarts.
27. The method of clause D16, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 5 quarts.
28. The method of clause D16, wherein the wherein the gearbox comprises a multi-stage planetary gearset.
29. A vertical take-off and landing (VTOL) aircraft comprising: at least four electrical propulsion systems, each electrical propulsion system comprising: an electrical motor, wherein the electrical motor includes at least a stator and a rotor; a gearbox assembly comprising: a planetary gear mechanically coupled to a rotor of an electric motor; a planetary carrier that is connected to at least one shaft from a set of shafts that extends concentrically from the at least one planetary gear; a main shaft comprising: a first end of the main shaft extending through the planetary carrier, a second end of the main shaft mechanically coupled to a propeller assembly, a length of the main shaft extending the first end of the main shaft through a sun gear and rotor to a second end of the main shaft; and a carrier cover that is connected to the main shaft and is connected to at least one shaft from a set of shafts that extends concentrically from the at least one planetary gear, wherein the electrical motor and gearbox assembly are concentrically aligned along the main shaft.
30. The electric propulsion system of clause D29, wherein a rotation of the carrier cover drives a rotation of the main shaft.
31. The electric propulsion system of clause D29, wherein the sun gear mechanically coupled to the rotor of the electric motor, wherein the electrical motor further comprises a stator.
32. The electric propulsion system of clause D29, wherein the sun gear is hollow.
33. The electric propulsion system of clause D29, further comprising a ring gear that interfaces with the at least one planetary gear.
34. The electric propulsion system of clause D29, wherein the planetary gear comprises a compound planetary gear.
35. The electric propulsion system of clause D29, wherein the gearbox comprises a multi-stage planetary gearset.
36. The electric propulsion system of clause D29, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 1 quart.
37. The electric propulsion system of clause D29, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 1.5 quarts.
38. The electric propulsion system of clause D29, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 2 quarts.
39. The electric propulsion system of clause D29, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 2.5 quarts.
40. The electric propulsion system of clause D29, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 3 quarts.
41. The electric propulsion system of clause D29, further comprising a cooling system that cools the gearbox with an amount of oil equal to or less than 5 quarts.
42. The electric propulsion system of clause D29, further comprising a gear ratio that is about 6.45.
43. The electric propulsion system of clause D29, wherein the sun gear, the planetary gear, the ring gear, the planetary carrier, and the carrier cover are disposed concentric with the main shaft.
44. The electric propulsion system of clause D29, wherein the electric motor and gearbox assembly are substantially aligned along the main shaft.
45. An electrical propulsion system having a gearbox apparatus that delivers power from an electrical engine via a reverse torque path comprising the electrical propulsion system of any clauses D1-D15 and D29-D45.
46. A method of delivering power from an electrical engine using a gearbox via a reverse torque path comprising the methods of any clauses D16-D28.
The embodiments disclosed herein are intended to be non-limiting. Those of ordinary skill in the art will appreciate that certain components and configurations of components may be modified without departing from the scope of the disclosed embodiments.
This disclosure claims priority to U.S. Provisional Application No. 63/378,536, titled “Tilt Rotor Systems and Methods for eVTOL Aircraft,” filed Oct. 6, 2022, and U.S. Provisional Application No. 63/378,680, titled “Systems and Methods for Improved Propulsion Systems for eVTOL Aircraft,” filed Oct. 7, 2022, the contents of which are incorporated herein in their entirety.
| Number | Name | Date | Kind |
|---|---|---|---|
| 2135477 | Griswold | Nov 1938 | A |
| 2477024 | Webster | Jul 1949 | A |
| 3087078 | Brown | Apr 1963 | A |
| 3217193 | Rayner | Nov 1965 | A |
| 3802795 | Nyeste et al. | Apr 1974 | A |
| 3909584 | Brienza et al. | Sep 1975 | A |
| 4092946 | Kappas | Jun 1978 | A |
| 4544166 | Karasawa | Oct 1985 | A |
| 5443130 | Tanaka et al. | Aug 1995 | A |
| 5461636 | Karube et al. | Oct 1995 | A |
| 5742484 | Gillette et al. | Apr 1998 | A |
| 5831358 | Bobay | Nov 1998 | A |
| 5833566 | Showalter | Nov 1998 | A |
| 6743135 | Klemen et al. | Jun 2004 | B2 |
| 7753151 | Marsh et al. | Jul 2010 | B2 |
| 8251861 | Varela | Aug 2012 | B2 |
| 8313403 | Wilson, Jr. | Nov 2012 | B2 |
| 8360928 | Tabata et al. | Jan 2013 | B2 |
| 8371978 | Nobata et al. | Feb 2013 | B2 |
| 8393791 | Thornton et al. | Mar 2013 | B2 |
| 8427019 | Garriga et al. | Apr 2013 | B2 |
| 8480528 | Wilhelm et al. | Jul 2013 | B2 |
| 8531072 | Wishart | Sep 2013 | B2 |
| 8649925 | Light et al. | Feb 2014 | B2 |
| 8650980 | Lafer et al. | Feb 2014 | B2 |
| 8692425 | Creviston et al. | Apr 2014 | B2 |
| 8771123 | Palfai et al. | Jul 2014 | B2 |
| 9022890 | Smetana | May 2015 | B2 |
| 9163715 | Valente | Oct 2015 | B2 |
| 9188195 | Iwasa | Nov 2015 | B2 |
| 9216804 | Van Der Laan et al. | Dec 2015 | B2 |
| 9272777 | McGlaun et al. | Mar 2016 | B2 |
| 9306433 | Sten et al. | Apr 2016 | B2 |
| 9413208 | Pal | Aug 2016 | B2 |
| 9482269 | Fruin et al. | Nov 2016 | B2 |
| 9677663 | Sammataro et al. | Jun 2017 | B2 |
| 9689466 | Kanai et al. | Jun 2017 | B2 |
| 9702255 | Martin et al. | Jul 2017 | B2 |
| 9787164 | Buttner et al. | Oct 2017 | B2 |
| 9917488 | Major | Mar 2018 | B2 |
| 9994305 | Moldovan | Jun 2018 | B1 |
| 10017247 | Elliott et al. | Jul 2018 | B1 |
| 10116187 | Wishart | Oct 2018 | B1 |
| 10145753 | Perelman et al. | Dec 2018 | B2 |
| 10220698 | Shibata et al. | Mar 2019 | B2 |
| 10302185 | Kawakami et al. | May 2019 | B2 |
| 10351235 | Karem et al. | Jul 2019 | B2 |
| 10392089 | Bentgen et al. | Aug 2019 | B2 |
| 10476351 | Francis et al. | Nov 2019 | B2 |
| 10502312 | Jaszewski et al. | Dec 2019 | B1 |
| 10598269 | Pydin | Mar 2020 | B2 |
| 10679586 | Poster | Jun 2020 | B2 |
| 10850841 | Bonden et al. | Dec 2020 | B2 |
| 10967934 | Ferguson | Apr 2021 | B2 |
| 11005318 | Nakane et al. | May 2021 | B2 |
| 11059598 | Arkus et al. | Jul 2021 | B2 |
| 11105340 | Cheung et al. | Aug 2021 | B2 |
| 11213934 | Bandy et al. | Jan 2022 | B2 |
| 11305869 | Ward | Aug 2022 | B1 |
| 11401042 | Peleg | Aug 2022 | B2 |
| 11401985 | Mantovan | Aug 2022 | B2 |
| 11413946 | Absenger et al. | Aug 2022 | B2 |
| 11414198 | Harvey | Aug 2022 | B2 |
| 11440401 | Sugimoto et al. | Sep 2022 | B2 |
| 11447004 | Misu et al. | Sep 2022 | B2 |
| 11453286 | Wishart | Sep 2022 | B2 |
| 11465488 | Li et al. | Oct 2022 | B2 |
| 20040195044 | Kimura et al. | Oct 2004 | A1 |
| 20050206251 | Foster | Sep 2005 | A1 |
| 20080022981 | Keyaki et al. | Jan 2008 | A1 |
| 20090206709 | Kakuda et al. | Aug 2009 | A1 |
| 20120021621 | Kim | Jan 2012 | A1 |
| 20120091833 | Hackett et al. | Apr 2012 | A1 |
| 20130126669 | Hamann et al. | May 2013 | A1 |
| 20130255416 | Terauchi | Oct 2013 | A1 |
| 20140151495 | Kuhn, Jr. | Jul 2014 | A1 |
| 20140274528 | Valente et al. | Sep 2014 | A1 |
| 20150044075 | Abe et al. | Feb 2015 | A1 |
| 20160013704 | Wagner et al. | Jan 2016 | A1 |
| 20160123417 | Rehfus et al. | May 2016 | A1 |
| 20160341281 | Brunvold et al. | Nov 2016 | A1 |
| 20180056774 | Chan | Mar 2018 | A1 |
| 20180180164 | Hagino | Jun 2018 | A1 |
| 20180283524 | Uematsu et al. | Oct 2018 | A1 |
| 20180363762 | Kiyokami et al. | Dec 2018 | A1 |
| 20190063503 | Nowoisky et al. | Feb 2019 | A1 |
| 20190211841 | Cottrell et al. | Jul 2019 | A1 |
| 20190285168 | Ikeda et al. | Sep 2019 | A1 |
| 20190319551 | Song et al. | Oct 2019 | A1 |
| 20190366951 | Young et al. | Dec 2019 | A1 |
| 20200140049 | Gil et al. | May 2020 | A1 |
| 20200149534 | Wetterau | May 2020 | A1 |
| 20200215920 | Suyama et al. | Jul 2020 | A1 |
| 20200262574 | Peleg | Aug 2020 | A1 |
| 20200361622 | Groninga et al. | Nov 2020 | A1 |
| 20210018089 | Guarino | Jan 2021 | A1 |
| 20210053433 | Gyarmati et al. | Feb 2021 | A1 |
| 20210104935 | Morrison | Apr 2021 | A1 |
| 20210107620 | Weekes et al. | Apr 2021 | A1 |
| 20210190201 | Shimokobe et al. | Jun 2021 | A1 |
| 20210226518 | Hung et al. | Jul 2021 | A1 |
| 20210229549 | Varela | Jul 2021 | A1 |
| 20210254705 | Olson et al. | Aug 2021 | A1 |
| 20210276707 | Heironimus et al. | Sep 2021 | A1 |
| 20210291971 | Bernard | Sep 2021 | A1 |
| 20210328482 | Rutowski et al. | Oct 2021 | A1 |
| 20210336510 | Yang et al. | Oct 2021 | A1 |
| 20210339881 | Bevirt et al. | Nov 2021 | A1 |
| 20210344255 | Said et al. | Nov 2021 | A1 |
| 20210403155 | Neiser | Dec 2021 | A1 |
| 20220069664 | Kim et al. | Mar 2022 | A1 |
| 20220119121 | Lacaux et al. | Apr 2022 | A1 |
| 20220135179 | Lee | May 2022 | A1 |
| 20220145806 | Beck et al. | May 2022 | A1 |
| 20220186670 | Blake et al. | Jun 2022 | A1 |
| 20220190764 | Yoon | Jun 2022 | A1 |
| 20220219817 | Knoll | Jul 2022 | A1 |
| 20220221048 | Remboski et al. | Jul 2022 | A1 |
| 20220250756 | Wagner et al. | Aug 2022 | A1 |
| 20220252146 | Hart et al. | Aug 2022 | A1 |
| 20220263384 | Kapatral et al. | Aug 2022 | A1 |
| 20220267020 | Warbeck | Aug 2022 | A1 |
| 20220289371 | Schroeder et al. | Sep 2022 | A1 |
| 20220294295 | Yang et al. | Sep 2022 | A1 |
| 20220297848 | Sivalingam et al. | Sep 2022 | A1 |
| 20220306300 | Matsumoto | Sep 2022 | A1 |
| 20220306305 | Cottrell et al. | Sep 2022 | A1 |
| 20220315237 | Fukuchi | Oct 2022 | A1 |
| 20220316582 | Inoue | Oct 2022 | A1 |
| 20220320957 | Tangudu | Oct 2022 | A1 |
| 20220360133 | Striedelmeyer | Oct 2022 | A1 |
| 20220357223 | Khazaai et al. | Nov 2022 | A1 |
| Number | Date | Country |
|---|---|---|
| 2507177 | Aug 2002 | CN |
| 201887602 | Jun 2011 | CN |
| 202997950 | Jun 2013 | CN |
| 206305855 | Jul 2017 | CN |
| 108390577 | Aug 2018 | CN |
| 208907678 | May 2019 | CN |
| 209516989 | Oct 2019 | CN |
| 113410060 | Apr 2022 | CN |
| 114583886 | Jun 2022 | CN |
| 202012101392 | May 2012 | DE |
| 102019213857 | Mar 2021 | DE |
| 0052015 | May 1982 | EP |
| 0488434 | Jun 1992 | EP |
| 2650567 | Oct 2013 | EP |
| 3139044 | Mar 2017 | EP |
| 2735501 | Jun 2018 | EP |
| 3078534 | Aug 2019 | EP |
| 3919376 | Dec 2021 | EP |
| 3939890 | Jan 2022 | EP |
| 4071983 | Oct 2022 | EP |
| 2012042806 | Apr 2012 | WO |
| 2014114560 | Jul 2014 | WO |
| WO-2017145833 | Aug 2017 | WO |
| 2017166811 | Oct 2017 | WO |
| 2018026735 | Feb 2018 | WO |
| 2019099378 | May 2019 | WO |
| 2021114606 | Jun 2021 | WO |
| 2021222528 | Nov 2021 | WO |
| 2022074405 | Apr 2022 | WO |
| 2022076531 | Apr 2022 | WO |
| 2022123229 | Jun 2022 | WO |
| 2022135939 | Jun 2022 | WO |
| 2022156546 | Jul 2022 | WO |
| 2022197838 | Sep 2022 | WO |
| 2022213209 | Oct 2022 | WO |
| 2022223810 | Oct 2022 | WO |
| Entry |
|---|
| “Flex without the Flex”, cy384, Nov. 12, 2020, http://www.cy384.com/blog/flex-pcbs.html. 5 pages. |
| “Endless-Snake Flexible Printed Circuits”, Leiton, May 20, 2022. 4 pages. |
| “Things to Know About PCB for LED Strips”, MyLedy, Sep. 13, 2021. 12 pages. |
| Ing, W. et al., “New Generation of Intelligent Electromechanical Valve Actuation”, 11th International Fluid Power Conference 2018, 5 pages. |
| Jeon, D. et al., “4 Types of Enclosures for Large Motors”, TMEIC, Jul. 4, 2021, 8 pages. |
| Number | Date | Country | |
|---|---|---|---|
| 63378680 | Oct 2022 | US | |
| 63378536 | Oct 2022 | US |
