System and method for electrical damping of aircraft flight control surfaces and propellers

Abstract

An aircraft system for damping the motion of structural flight elements. The aircraft system includes a movable aircraft structure, a power supply, a controller, a resistor, and a first transistor connected in series with the resistor. When the aircraft power supply is providing power, the controller controls the first transistor to prevent current from flowing through its terminals and the resistor. Furthermore, when the aircraft power supply is not providing power, the first transistor allows current generated by a reverse EMF (electromotive force) voltage to flow through its terminals and the resistor, and the reverse EMF voltage is generated by the motion of the movable aircraft structure.

Description

Technical Field

[0001] Cross - Reference to Related Applications This application claims the priority and benefit of U.S. Non - Provisional Application No. 18 / 401,446 (Attorney Docket No. 16163.0026 - 00000) entitled "Systems and Methods for Electric Damping of Aircraft Flight Control Surfaces and Propellers" filed on December 30, 2023, which in turn claims the priority and benefit of U.S. Provisional Application No. 63 / 459,324 (Attorney Docket No. 16163.6010 - 00000) entitled "Systems and Methods for Electric Damping of Aircraft Flight Control Surfaces and Propellers" filed on April 14, 2023. The entire content of the above - mentioned applications is incorporated herein in its entirety for all purposes.

[0002] The present disclosure generally relates to the field of powered aircraft. More particularly, without limitation, the present disclosure relates to innovations in tilt - rotor aircraft using electric propulsion systems. Certain aspects of the present disclosure generally relate to electrically damping motor - controlled aircraft flight elements. Further, certain aspects of the present disclosure relate to motor control systems and associated circuitry that use the reverse EMF (electromotive force) voltage created by a moving flight element to generate a counter - torque against the movement.

Summary of the Invention

[0003] Embodiments of the present disclosure provide a system for damping the movement of an aircraft flight element. As used herein, an aircraft flight element can refer to one or more movable structural components of an aircraft, such as flight control surfaces (e.g., ailerons, elevators, rudders, etc.), rotors, prop - rotors, and / or propellers.

[0004] One aspect of the present disclosure relates to an electrical system for an aircraft, comprising a controller, a resistor, and a switching device connected in series with a first resistor. The controller is configured to control the switching device to prevent current from flowing through the first resistor when the aircraft's power supply is providing power. The switching device is configured to allow current generated by a reverse EMF (electromotive force) voltage to flow through the first resistor when the aircraft's power supply is not providing power, the reverse EMF voltage is generated by the movement of the aircraft's movable aircraft structures. [Brief explanation of the drawing]

[0005] [Figure 1A] An exemplary VTOL aircraft in a cruising configuration consistent with embodiments of this disclosure is illustrated.

[0006] [Figure 1B] An exemplary VTOL aircraft in an ascent configuration consistent with embodiments of this disclosure is illustrated.

[0007] [Figure 2] A motor control system including a brake resistor in a loaded state, consistent with embodiments of this disclosure, is illustrated as an example.

[0008] [Figure 3] A motor control system including a no-load braking resistor, consistent with embodiments of this disclosure, is illustrated as an example.

[0009] [Figure 4] A motor control system including a damping resistor between two depletion-mode transistors, consistent with embodiments of this disclosure, is illustrated.

[0010] [Figure 5] A motor control system including relays, consistent with embodiments of this disclosure, is illustrated as an example.

[0011] [Figure 6]A motor control system including a set of depletion-mode transistors consistent with embodiments of this disclosure is illustrated.

[0012] [Figure 7A] A motor control system including a passive discharge resistor and a diode in a capacitor discharge state, consistent with embodiments of this disclosure, is illustrated.

[0013] [Figure 7B] A motor control system including a passive discharge resistor and a diode in a decayed state, consistent with embodiments of this disclosure, is illustrated as an example.

[0014] [Figure 8] A method for mitigating short-circuit failures consistent with the embodiments of this disclosure is illustrated below. [Modes for carrying out the invention]

[0015] This disclosure addresses components of an electric vertical take-off and landing (eVTOL) aircraft, primarily for use in non-conventional aircraft. For example, the eVTOL aircraft of this disclosure may be intended for frequent (e.g., more than 50 flights per working day), short-duration flights (e.g., less than 100 miles per flight) over, into, and outside densely populated areas. The aircraft may be intended to carry 4 to 6 passengers or commuters who expect a low-noise and low-vibration experience. Therefore, it may be desirable that their components be configured and designed to withstand frequent use without wear, generate little heat and vibration, and include mechanisms that allow the aircraft to effectively control and manage the heat or vibration generated by the components. Furthermore, some of these aircraft may be intended to operate in close proximity to one another over congested metropolitan areas. Therefore, it may be desirable that their components be configured and designed to generate low levels of noise inside and outside the aircraft and have various safety and backup mechanisms. For example, for safety reasons, it may be desirable for an aircraft to be propelled by a distributed propulsion system to avoid the risk of a single point of failure and to be able to perform conventional takeoffs and landings on a runway. Furthermore, it may be desirable for an aircraft to be able to safely take off and land vertically from a relatively limited space (e.g., a vertiport, parking lot, or private road) compared to a conventional airport runway, while transporting approximately 4-6 passengers or commuters along with their luggage. These usage requirements may impose design constraints on the size, weight, and operational efficiency (e.g., drag, energy use) of the aircraft, which may affect the design and configuration of aircraft components.

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

[0017] In some embodiments, the eVTOL aircraft of this disclosure may be designed to be capable of both vertical and conventional takeoffs and landings, with a distributed electric propulsion system that enables vertical flight, forward flight, and transition. Thrust may be generated by supplying high-voltage power to electric engines of the distributed electric propulsion system, each capable of converting the high-voltage power into mechanical shaft power to rotate a propeller. Embodiments disclosed herein may involve optimizing the energy density of the electric propulsion system. Embodiments may include electric engines connected to an onboard power supply, which may include a device capable of storing energy such as a battery or capacitor, or one or more systems for utilizing or generating electricity such as a fuel generator or a solar panel array. Some disclosed embodiments provide weight reduction and space saving of components within the aircraft, thereby improving the efficiency and performance of the aircraft. Focusing on safety in passenger transport, the disclosed embodiments implement new and improved safety protocols and system redundancy in case of failure to minimize any single point of failure in the aircraft propulsion system. Some disclosed embodiments also provide new and improved approaches to meeting aviation and transport laws and regulations.

[0018] In a preferred embodiment, the distributed electric propulsion system may include twelve electric engines that can be mounted on forward and aft booms of the aircraft's wings. The forward electric engines may be able to tilt during flight between a horizontally oriented position (e.g., to generate forward thrust) and a vertically oriented position (e.g., to generate vertical lift). The forward electric engines may be of a clockwise or counterclockwise type with respect to the direction of propeller rotation. The aft electric engines may be fixed in a vertically oriented position (e.g., to generate vertical lift). They may also be of a clockwise or counterclockwise type with respect to the direction of propeller rotation. In some embodiments, the aircraft may have various combinations of forward and aft electric engines. For example, the aircraft may have six forward electric engines and six aft electric engines, four forward electric engines and four aft electric engines, or any other combination of forward and aft engines, including embodiments in which the number of forward and aft electric engines are not equal. In some embodiments, the aircraft may have four forward propellers and four rear propellers, at least four of which include tiltable propellers.

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

[0020] In some embodiments, in a conventional conventional takeoff and landing (CTOL) mission, the forward electric engine may provide horizontal thrust for wing-supported takeoff, cruise, and landing. In some embodiments, the rear electric engine may not be used to generate thrust during the CTOL mission, and the rear propeller may be housed in a predetermined position.

[0021] In some embodiments, the electric engine is housed or connected to a boom of the aircraft and may include a motor, an inverter, and a gearbox. In some embodiments, the motor, the inverter, and the gearbox may be interfaced so that they share a central axis. In some embodiments, the torque resulting from the motor may be sent to the gearbox away from the propeller of the propulsion system. In some embodiments, the gearbox may provide gear reduction and then send the torque back to the propeller through a bearing located inside the motor via a main shaft. In some embodiments, the inverter may be mounted at the rear of the gearbox so that the main shaft does not move through the inverter when outputting torque to the propeller.

[0022] In some embodiments, the tilt propeller system may include a linear or rotary actuator for changing the orientation of the propulsion system during operation. In some embodiments, the pitch of the propulsion system may be changed as a function of the orientation of the propulsion system. In some embodiments, the rotary actuator may include a motor, an inverter, and a gearbox. In some embodiments, the gearbox may include various types of gears that interface to provide a gear reduction that enables the propulsion system to be oriented. In some embodiments, the tilt propeller system may include a redundant configuration in which there are multiple motors, inverters, and gearboxes and the gears are used to interface. In some embodiments, a configuration that utilizes multiple motors, gearboxes, and inverters may allow a failed portion of the redundant configuration to be driven by motors, inverters, and gearboxes of another portion of the configuration. In some embodiments, the gearbox configuration may also allow the tilt propeller system to maintain the orientation of the propulsion system with the aid of additional power provided by the system or without additional power.

[0023] In some embodiments, the electric propulsion systems described herein may generate thrust by supplying high voltage (HV) power to an electric engine, which then converts the HV power into mechanical shaft power used to rotate a propeller. As described above, the aircraft described herein may have multiple electric engines mounted on booms in front of and behind the wings. The amount of thrust generated by each electric engine may be controlled by a torque command from a flight control system (FCS) via a digital communication interface to each electric engine. Embodiments may include a forward electric engine, the orientation or tilt of which may be changed. Additional embodiments may include a forward engine that may be of a clockwise (CW) type or a counterclockwise (CCW) type. The forward electric engine propulsion subsystem may be composed of a multi-blade adjustable pitch propeller, as well as a variable pitch subsystem.

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

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

[0026] As disclosed herein, an electric engine may include an inverter and motor, or an inverter, gearbox, and motor, across a variety of configurations, such as the representative configuration described herein. For example, an electric engine may include an electric motor, gearbox, and inverter, all sharing the same central axis. In addition, the central axis may be configured along the axis of an output shaft directed toward the aircraft's propeller. In such an exemplary configuration, the motor, gearbox, and inverter would all share the output shaft as the central axis and be oriented circularly around the output shaft. Additional embodiments may include a motor, gearbox, and inverter mounted together in sequence, or a configuration in which some of the components, such as a motor and gearbox, are mounted together, and other components, such as an inverter, are located elsewhere, but a wiring system is used to connect the electric engine.

[0027] As described herein, the electric engines for aircraft described herein may include some or all of a motor, inverter, and gearbox. Various configurations may include an inverter and motor such that the motor's output shaft directly provides speed and torque to the propeller shaft. Additional embodiments of the electric engine may include a motor, inverter, and gearbox, where the motor's output is transmitted through a gearbox connected to an output shaft for the propeller, or where the motor, inverter, and gearbox are combined, with the motor's output traveling away from the propeller and through the gearbox, and the output shaft for the propeller returning to the propeller via the gearbox and motor. As described herein, the electric engine may consider any combination or orientation of some or all of the motor, inverter, and gearbox. In addition, each configuration or orientation of the electric engine disclosed herein may include cooling via air cooling, coolant, or a mixture of both.

[0028] For example, the configuration of an electric engine may include a motor and an inverter, with the motor located between the aircraft's propeller and the inverter. In addition, the motor may include a gearbox. Furthermore, the inverter may share the same central axis as the motor and may be located in a cantilevered housing supported from the rear of the motor, and may be air-cooled. It is recognized that such an orientation of the inverter may not be the optimal configuration with respect to the housing required to realize such a cantilevered orientation. In addition, the motor in this configuration utilizing air cooling may be equipped with potting material and air fins to assist in cooling the motor, which may lead to a further significant increase in the system's mass.

[0029] Some embodiments may include an electric engine in which the inverter module may be mounted on the outside of the motor housing. Additional embodiments may include an electric engine in which the inverter may be mounted on top of the electric motor such that the cooling fins of the inverter are located below the propeller.

[0030] Embodiments of an electric motor may include a stator housing, a wound stator assembly, a rotor, various bearings, and optional additional components to assist in transmitting the speed and torque generated by the motor to a propeller.

[0031] It is understood that electric engines can generate heat during operation and may be equipped with a thermal management system to ensure that the components of the electric engine do not fail during operation. In some embodiments, a coolant may be used and circulated throughout the individual components of the engine, such as an inverter, gearbox, or motor, through some of the components, or through all of the components of the engine, to assist in managing the heat present in the engine. Additional embodiments may include using an air cooling method to cool the electric engine, or using a mixture of coolant and air to manage the heat generated during the operation of the electric engine. In some embodiments, the coolant used may also be the same liquid used as a lubricant throughout the inverter, gearbox, or motor. For example, the inverter, gearbox, and motor may be cooled using liquid or air, or a mixture of air and liquid cooling may be used, such as using air cooling to cool the motor and using liquid cooling in the inverter and gearbox, or any other combination of air and liquid cooling may be used over the inverter, gearbox, and motor, or even a subset of their components.

[0032] In some embodiments, oil may be used as a lubricant throughout the electric engine and may also be used as a coolant fluid to help manage the heat generated by the engine during operation. In addition to this example, different amounts of oil may be used to act as both a lubricant and a coolant fluid in the electric engine, either in combination with or without air cooling assistance, such as less than one quart, less than two quarts, or any other amount of oil necessary to lubricate and cool the electric engine. As disclosed herein, electric engines may have different primary functions, for example, used only for ascent and landing, and in that case used only in one direction, or used in all stages of flight, such as ascent, landing, and during flight.

[0033] Additional embodiments of the electric engine may include active protection features in the forward and rear electric engine, such as monitoring of internal temperatures throughout engine operation, including oil temperature, stator winding set, inverter bulk capacitor, power module, control panel power module, control panel control processor, control panel monitoring processor, internal hot spots, and various other locations throughout the engine. Embodiments may include over-temperature limits that take into account known fault temperatures and operating limits related to the autoignition temperature of the fluid. Some embodiments may include a high-voltage power system that may have fuses at the high-voltage battery terminals that can irreversibly and quickly disconnect the engine's electrical connections to mitigate overcurrent events. This overcurrent protection may be activated when the current draw of the electric engine is greater than the overcurrent operating point. Thus, in some embodiments, fault conditions leading to overcurrent may only lead to transient overheating, arcing, or sparking faults.

[0034] As described above, an aircraft embodiment may include many movable structural flight elements that enable the pilot to safely control the aircraft. Flight control surfaces (e.g., ailerons, elevators, rudders, etc.) are important for controlling the aircraft's positioning. Changes in the orientation of these surfaces alter the airflow and pressure distribution around the aircraft, allowing the pilot to control the aircraft's movement on three axes of rotation. Similarly, control of the rotation and orientation of rotors and prop rotors can provide the lift necessary for vertical takeoff, landing, and hovering. The rotation of propellers can provide the thrust necessary to move the aircraft in the air. The movement of each of these flight elements is important for the aircraft's safety and stability. In normal operation during flight, power motors control the movement of these flight elements.

[0035] When an aircraft is parked or taxiing on the ground, the power may be cut off, and the motors of these flight elements may be unloaded. On windy days, gusts of wind on the ground may cause uncontrolled movement of flight elements, potentially damaging them and creating a safety hazard for subsequent flights. Furthermore, this movement may create dangerous conditions for maintenance personnel.

[0036] Similarly, during flight, failures and / or loss of power to flight element motors can cause uncontrolled movement of flight elements. For example, failures and / or loss of power to flight control surfaces can contribute to aerodynamic flutter, leading to excessive movement and damage to the aircraft's flight elements. Damage to flight elements can be costly to repair, result in aircraft downtime, or, if undetected, lead to unsafe flight conditions. For example, failures, loss of engine power, and / or low propeller pitch angles can increase propeller rotation speed to levels exceeding their structural dynamic limits or the structural integrity of the hub or blades, leading to catastrophic aircraft conditions.

[0037] The Federal Aviation Administration (FAA) recognizes the risks associated with these movable flight elements and has issued guidelines for securing moving parts and limiting propeller speed during flight. However, securing procedures are not always feasible, such as when weather conditions make it dangerous to approach the aircraft or when the aircraft is in flight. Therefore, it is necessary to dampen the movement of movable flight elements to reduce or eliminate the possibility of damage to these elements.

[0038] Furthermore, resolving the problem of uncontrolled flight element movement is particularly important in electric aircraft or other aircraft that use electrically actuated flight elements. Conventional gas-powered aircraft typically include hydraulically actuated flight elements, where residual hydraulic pressure acts to dampen the movement of the flight elements. When flight elements are electrically actuated, there is no hydraulic pressure to help dampen their movement. Some aircraft with electrically actuated flight elements employ mass balance devices to help dampen the movement of the flight elements. However, these mass balance devices add additional weight to the aircraft, negatively impacting its range and maneuverability. Therefore, it is necessary to dampen the movement of electrically actuated flight elements.

[0039] Embodiments of this disclosure solve these and other problems by generating a counter-torque to counteract the motion of the flight element and by dissipating excess energy as heat. As a result, even when the motor associated with the flight element is unloaded, the motor control system dampens the motion of the flight element and avoids damage to the aircraft and / or dangerous conditions.

[0040] Herein, exemplary embodiments are given in detail, illustrated in the accompanying drawings. The following description is given with reference to the accompanying drawings, and unless otherwise specified, the same reference numerals in different drawings represent the same or similar elements. The implementations described below in the description of exemplary embodiments are not representative of all implementations consistent with this disclosure. Rather, they are merely examples of apparatus and methods consistent with aspects relating to the subject matter described in the accompanying claims.

[0041] Figures 1A and 1B illustrate VTOL aircraft 100 in a cruising configuration and a vertical takeoff, landing, and hovering configuration (also referred to herein as a “climb” configuration), respectively, consistent with embodiments of the present disclosure. The aircraft 100 may include a fuselage 102, wings 104 mounted on the fuselage 102, a tail 105, and one or more rear stabilizers 106 mounted on the tail 105 or the rear of the fuselage 102. Multiple lift propellers 112 may be mounted on the wings 104 and may be configured to provide lift for vertical takeoff, landing, and hovering. Multiple tilt propellers 114 may be mounted on the wings 104 and may be tiltable between a cruising configuration, as shown in Figure 1A, in which the multiple tilt propellers 114 provide forward thrust to the aircraft 100 for horizontal flight, and an ascent configuration, as shown in Figure 1B, in which the multiple tilt propellers 114 provide some of the lift necessary for vertical takeoff, landing, and hovering. As used herein, the ascent configuration may refer to an orientation of the tilt propellers in which the thrust of the tilt propellers continues to provide primarily lift to the aircraft. The cruising configuration may refer to an orientation of the tilt propellers in which the thrust of the tilt propellers continues to provide primarily forward thrust to the aircraft. Alternatively, the cruising configuration may refer to a configuration in which a lift propeller is housed.

[0042] In some embodiments, the lift propeller 112 may be configured to provide only lift, with all thrust provided by the tilt propeller. Thus, the lift propeller 112 may be in a fixed position and may generate thrust only during takeoff, landing, and hovering. On the other hand, the tilt propeller 114 may be tilted into an upward configuration so that their thrust is directed downward to provide additional lift.

[0043] For forward flight, the tilt propellers 114 can be tilted from their climb configuration to their cruising configuration. In other words, the pitch and tilt angles of the tilt propellers 114 can be changed from a direction in which the tilt propeller thrust is directed downward (to provide lift during vertical takeoff, landing, and hovering) to a direction in which the tilt propeller thrust is directed aft (to provide forward thrust to the aircraft 100). The tilt propellers can be tilted around an axis that may be perpendicular to the forward direction of the aircraft 100. When the aircraft 100 is in full forward flight in the cruising configuration, lift can be fully provided by the wings 104. The lift propellers 112, on the other hand, can be stopped. The blades 120 of the lift propellers 112 can be locked in a low-drag position for aircraft cruising. In some embodiments, each lift propeller 112 may have two blades 120 that can be locked to cruise in a minimum drag position where one blade is directly in front of the other, as illustrated in Figure 1A. In some embodiments, the lift propeller 112 has three or more blades. In some embodiments, the tilt propeller 114 includes more blades 118 than the lift propeller 112. For example, as illustrated in Figures 1A-B, each lift propeller 112 may include, for example, two blades, and each tilt propeller 114 may include, for example, five blades. In some embodiments, the tilt propeller 114 may have, for example, two to five blades.

[0044] In some embodiments, the aircraft may include only one wing 104 on each side of the fuselage 102 (or a single wing extending over the entire aircraft), with at least a portion of the lift propellers 112 located behind the wing 104 and at least a portion of the tilt propellers 114 located in front of the wing 104. In some embodiments, all of the lift propellers 112 may be located behind the wing 104, and all of the tilt propellers 114 may be located in front of the wing 104. According to some embodiments, all of the lift propellers 112 and tilt propellers 114 may be mounted on the wing, i.e., the lift propellers or tilt propellers may not be mounted on the fuselage. In some embodiments, all of the lift propellers 112 may be located behind the wing 104, and all of the tilt propellers 114 may be located in front of the wing 104. According to some embodiments, all of the lift propellers 112 and tilt propellers 114 may be located inside the wingtip 109.

[0045] In some embodiments, the lift propeller 112 and the tilt propeller 114 may be mounted on the wing 104 by a boom 122. The boom 122 may be mounted below the wing 104, above the wing, and / or integrated into the wing profile. In some embodiments, one lift propeller 112 and one tilt propeller 114 may be mounted on each boom 122. The lift propeller 112 may be mounted at the rear end of the boom 122, and the tilt propeller 114 may be mounted at the front end of the boom 122. In some embodiments, the lift propeller 112 may be mounted in a fixed position on the boom 122. In some embodiments, the tilt propeller 114 may be mounted at the front end of the boom 122 via a hinge. The tilt propeller 114 can be mounted on the boom 122 such that, when in the cruising configuration, the tilt propeller 114 is aligned with the body of the boom 122, forming a continuous extension of the front end of the boom 122 that minimizes drag for forward flight.

[0046] In some embodiments, the aircraft 100 may include, for example, one wing on each side of the fuselage 102, or a single wing extending across the entire aircraft. According to some embodiments, at least one wing 104 is a high wing mounted on the upper side of the fuselage 102. According to some embodiments, the wing includes control surfaces such as flaps, ailerons, or flaperons. According to some embodiments, the wing may have curved wingtips 109 to reduce drag during forward flight.

[0047] In some embodiments, the rear stabilizer 106 includes control surfaces such as one or more rudders, one or more elevators, and / or one or more rudders and elevators combined. The (one or more) wings may have any preferred design. For example, the wings may have a tapered leading edge or a tapered trailing edge. In some embodiments, the wings may have a substantially straight leading edge in the central section of the wing 104.

[0048] Figure 2 illustrates a motor control system 200 operating under a load condition, consistent with embodiments of the present disclosure. The motor control system 200 may include a power supply 204. In some embodiments, the power supply 204 may include a single battery, while in other embodiments, the power supply 204 may include multiple batteries arranged in series, parallel, or a combination of series and parallel. In some embodiments, the power supply is a battery pack and battery management system, as detailed in PCT / US2023 / 79690, which is incorporated herein by reference in its entirety. In Figure 2, showing the “load” condition, the power supply 204 provides power to the circuit, as schematically illustrated by a closed switch 211. Current 212 flows from the negative side to the positive side of the power supply, providing power to the motor.

[0049] In some embodiments, the motor may be a permanent magnet motor, comprising one or more rotating permanent magnets connected to a gearbox and / or one or more flight elements. In some embodiments, the rotor may include permanent magnets, and stator windings (e.g., 210a-210c) may generate a magnetic field. In some embodiments, the motor may be a single-phase, two-phase, or three-phase motor, while in other embodiments, the motor may have a different number of phases. In some embodiments, the motor is a DC motor, such as a brushless DC motor. In some embodiments, the DC motor controls one or more flight control surfaces (e.g., slats, ailerons, tabs, spoilers, elevators, rudder, etc.), while in other embodiments, the DC motor controls one or more other flight elements.

[0050] In some embodiments, the motor is an AC motor. In an AC motor configuration, a controller 201 and / or a motor bridge 202 may provide a sinusoidal output voltage to stator windings 210a-210c. In some embodiments, the sinusoidal output voltage is produced by the controller 201 and / or the motor bridge 202, which switches transistors 208a-c and 209a-c on and off using pulse-width modulation (PWM). In some embodiments, the AC motor configuration controls a rotor, prop rotor, and / or propeller, while in other embodiments, the AC motor controls one or more other flight elements. In some embodiments, the motor may include three stator windings, as illustrated by 210a-210c.

[0051] In some embodiments, the capacitor 205 is connected in parallel with the power supply 204. In some embodiments, the capacitor 205 may be a bulk capacitor. The bulk capacitor 205 stores charge and ensures that a stable current supply is provided to the connected circuit. Furthermore, in some embodiments, the bulk capacitor 205 may provide a path for ripple current arising from a rotating magnetic field.

[0052] The controller 201 and / or motor bridge 202 may be hardware devices such as a computer, processor, and / or microprocessor. In some embodiments, the controller 201 and / or motor bridge 202 may receive control signal commands from a flight control system, flight control computer, autopilot, and / or other computer systems that regulate aircraft flight control. In some embodiments, the controller 201 is part of a flight control system, flight control computer, autopilot, and / or other computer systems that regulate aircraft flight control. In some embodiments, the controller 201 receives signals from a battery management system, battery management unit, and / or any other device configured to monitor and / or control the power supply 204. In some embodiments, the controller 201 and / or motor bridge 202 may be housed together with transistors, damping resistors, and / or other motor control elements. In other embodiments, the controller and / or motor bridge 202 may be housed separately from transistors, damping resistors, and / or other motor control elements.

[0053] Based on the received control signal commands, the controller 201 and / or motor bridge 202 may adjust the voltages to transistors 208a~c and 209a~c. In some embodiments, the controller 201 may provide signals to the motor bridge 202, while in other embodiments, the controller 201 may directly adjust the voltages to downstream devices such as transistors 208a~c and / or 209a~c. The motor Hall sensor 203 may provide the controller 201 with information regarding the positioning and speed of a magnetic rotor (not shown), which may be modified based on feedback.

[0054] Transistors 208a-c and 209a-c connect and disconnect the flow of current based on whether a voltage is received from the controller 201 and / or the motor bridge 202. In some embodiments, the transistors are bipolar transistors, field-effect transistors (e.g., MOSFETs), and / or insulated-gate bipolar transistors (e.g., IGBTs). In some embodiments, transistors 208a-c and 209a-c are enhancement-mode transistors, and when the controller 201 and / or the motor bridge 202 generates a voltage across the transistors (208a-c and / or 209a-c), the transistors allow current from the power supply 204 to propagate across their terminals, and the power can be connected to the stator windings 210a-210c. In other embodiments, transistors 208a-c and 209a-c are depletion-mode transistors, and when the controller 201 and / or motor bridge 202 do not generate voltage across the transistors (208a-c and / or 209a-c), the transistors allow current from the power supply 204 to propagate across their terminals, thereby connecting power to the stator windings 210a-210c. In some embodiments, transistors 208a-c are enhancement-mode transistors, and transistors 209a-c are depletion-mode transistors. In some embodiments, transistors 208a-c are depletion-mode transistors, and transistors 209a-c are enhancement-mode transistors. The controller 201 and / or motor bridge 202 selectively control the power to each stator winding 210a-210c to generate a magnetic field and control the associated flight elements.

[0055] The stator windings 210a to 210c are coils of conductors (e.g., insulated aluminum or copper wire) that each generate a magnetic field when energized. Figure 2 shows three stator windings, but a stator can contain any number of windings. For example, a motor may contain one, two, or any number of windings. The stator windings can also be wound in a variety of different ways. In some embodiments, the stator windings may be concentrated, while in other embodiments, the windings may be distributed.

[0056] As discussed above, the controller 201 and / or the motor bridge 202 can selectively control the voltage across each stator winding 210a to 210c to generate a magnetic field. For example, in the first motor phase, when stator windings 210c and 210a are magnetized, the controller 201 and / or the motor bridge 202 can control transistors 208c and 209a to allow current to flow through their terminals while keeping the remaining transistors closed. The current can flow through transistor 208c, through stator winding 210c, through stator winding 210a, and through transistor 209a to form a completed circuit. In the second motor phase, when stator windings 210a and 210b are magnetized, the controller 201 and / or the motor bridge 202 can control transistors 208b and 209a to allow current to flow through their terminals while keeping the remaining transistors closed. Current can flow through transistor 208b, through stator winding 210b, through stator winding 210a, and through transistor 209a, forming a completed circuit. In the third motor phase, when stator windings 210b and 210c are magnetized, controller 201 and / or motor bridge 202 may control transistors 208c and 209b to allow current to flow through their terminals, while keeping the remaining transistors closed. Figure 2 shows an exemplary embodiment, but different combinations of phases and / or windings may be used.

[0057] The magnetic fields created by the stator windings 210a-210c interact with a magnetic rotor (not shown), causing the rotor to rotate or otherwise move. The rotor can be operably coupled to a movable flight element. For example, the rotor can be operably coupled to a flight control surface (e.g., slats, ailerons, tabs, spoilers, elevators, rudder, etc.), a rotor, a prop rotor, and / or a propeller. Thus, under load conditions, the controller 201 and / or motor bridge 202 control the movement of the flight elements by adjusting the power from the power supply 204 to the stator windings 210a-210c. In some embodiments, the stator windings 210a-210c may be controlled to increase the speed of the propeller, rotor, and / or prop rotor. In some embodiments, the stator windings 210a-210c may be controlled to adjust the position (e.g., angle) of the flight control surface, or the direction of the tilt of the rotor and / or prop rotor.

[0058] The switching device connects and disconnects the flow of current across the circuit through the damping resistor 206. For example, the depletion-mode transistor 207 connects and disconnects the flow of current across the entire circuit through the damping resistor 206 based on whether a voltage is received from the controller 201 and / or the motor bridge 202. In some embodiments, the depletion-mode transistor 207 is a bipolar transistor, a field-effect transistor (e.g., a MOSFET), or an insulated-gate bipolar transistor (e.g., an IGBT). In a load condition with the power supply 204 connected, the controller 201 and / or the motor bridge 202 generate a voltage across the depletion-mode transistor 207, and the transistor does not allow current from the power supply 204 to propagate across its terminals. Thus, in a load condition, the depletion-mode transistor 207 keeps the damping resistor 206 disconnected from the rest of the circuit, as shown in Figure 2.

[0059] The damping resistor 206 is connected in series with the depletion-mode transistor 207 and is used to dissipate heat and slow down and / or stop the moving flight element. In some embodiments, the damping resistor may have a fixed resistance value, while in other embodiments, the damping resistor 206 may have an adjustable variable resistance value. In some embodiments, a resistance value that is the minimum resistance appropriate to dampen the motion of the flight element may be selected when considering the reverse EMF voltage of the motor. The resistance may be selected based on a function of resistance = (voltage)^2 / (damping force), where the damping force is based on the product of the hinge moment and velocity of the flight element. A smaller resistance value may be used when a larger reverse torque is required to counteract the motion of the flight element. A smaller resistance value allows a larger current flow through the stator windings 210a~210c, which can create a stronger magnetic field to counteract the motion of the rotor and connected flight element. In contrast, a larger resistance value may be used when a smaller reverse torque is required to counteract the motion of the flight element. A higher resistance value allows a smaller current to flow through the stator windings 210a-210c, potentially creating a weaker magnetic field to counteract the motion of the rotor and connected flight elements. In some embodiments, the damping resistor 206 may be a single resistor, while in other embodiments, the damping resistor 206 may include multiple resistors arranged in series, parallel, or a combination of series and parallel.

[0060] In some embodiments, the damping resistor 206 is capable of dissipating the large amount of heat generated by the current produced. In some embodiments, the damping resistor 206 may be coupled with a heat sink and / or heat exchanger that allows heat to be transferred away from the resistor. For example, the resistor and / or other circuit elements may be connected to a folded-fin heat exchanger or any other type of heat exchanger. In some embodiments, oil or other coolant can be circulated to remove heat from the damping resistor 206. In some embodiments, the placement of the damping resistor 206 on the circuit board may allow heat dissipation away from other board components.

[0061] Figure 3 illustrates a motor control system 200 operating in a no-load state with power supply 204 disconnected, consistent with embodiments of the present disclosure. Power supply 204 may be disconnected because the aircraft is parked on the ground, experiencing a power outage, experiencing an electrical circuit problem, experiencing a flight element problem, or for any other reason. In some embodiments, power supply 204 may be disconnected based on a mode selected by the aircraft pilot (e.g., off mode, standby mode, ground mode, etc.). In some embodiments, power supply 204 may be disconnected based on the detection of a short circuit and / or overcurrent condition. For example, a battery management system associated with power supply 204 and / or controller 201 may detect a short circuit and / or overcurrent condition and disconnect power supply 204 (e.g., by blowing a fuse).

[0062] In some embodiments, the power supply 204 may be shut off based on the detection of problems relating to the positioning and / or movement of the flight elements. For example, the flight control system and / or controller 201 may determine that the positioning of the flight elements is incorrect (e.g., the propeller pitch is too low) and / or that the movement of the flight elements is incorrect or uncontrolled. Based on this determination, the flight control system, battery management system, and / or controller 201 may shut off the power supply 204.

[0063] Based on the movement of air around the aircraft, the flight elements and mounted magnetic rotors may begin to move relative to the stator. The rotating magnetic field created by the rotor movement interacts with the stator coils 210a-210c, generating a current throughout the circuit, as indicated by a counterclockwise current flow 212. In some embodiments, during no-load conditions, the controller 201 and / or motor bridge 202 may receive a signal to provide voltage to the enhancement-mode transistors (208a-208c and 209a-209c), allowing the generated current to propagate through their terminals into the circuit. In other embodiments, the transistors (208a-208c and 209a-209c) may be depletion-mode transistors, and if they do not receive voltage from the controller 201 and / or motor bridge 202, the transistors may allow the generated current to propagate through their terminals into the circuit. Similarly, under no-load conditions, the controller 201 no longer supplies voltage to the depletion-mode transistor 207, allowing the generated current to pass across its terminals and through the damping resistor 206.

[0064] The current flow through the connected circuit and damping resistor 206 generates a magnetic field in the stator windings 210a-210c that is related to the rotational speed of the rotating magnetic rotor (not shown). These magnetic fields interact with the rotating magnetic rotor (not shown) and can provide a counter-torque that opposes the rotation, with stronger counter-torque provided at higher rotor speeds. Thus, the connected circuit and damping resistor 206 dampen the motion of the rotor and connected flight elements. Furthermore, the damping resistor 206 allows excess energy to be dissipated in the form of heat.

[0065] Figure 4 illustrates a motor control system 400 that includes a damping resistor 206 between two depletion-mode transistors (207 and 401) consistent with embodiments of the present disclosure. Figure 4 includes the same devices as the control system 200 in Figures 2 and 3. As will be further detailed below, the switching device further comprises a second depletion-mode transistor 401. In some embodiments, the depletion-mode transistor 401 is a bipolar transistor, a field-effect transistor (e.g., a MOSFET), or an insulated-gate bipolar transistor (e.g., an IGBT). Similar to the depletion-mode transistor 207, the depletion-mode transistor 401 connects and disconnects the flow of current across the circuit through the damping resistor 206 based on whether a voltage is received from the controller 201. Under load conditions, the controller 201 generates a voltage across the depletion-mode transistor 401, and the transistor does not allow current from the power supply 204 to propagate across its terminals. Therefore, the damping resistor 206 remains disconnected from the rest of the circuit when the power supply 204 is connected. However, as with the motor control system 200, in the no-load state, the controller 201 no longer supplies voltage to the depletion-mode transistors 207 and 401, allowing the generated current to pass through their terminals and through the damping resistor 206.

[0066] By providing two depletion-mode resistors (207 and 401), the control system 300 provides redundancy, avoids single-point failures, and increases the reliability of motor control. It is important that the braking resistor 206 remains disconnected during load operation (e.g., during flight) with the power supply 204 connected. Connecting the braking resistor 206 under load conditions could cause the resistor 206 to overheat, resulting in power consumption and damage to the resistor 206, the circuit, and / or the stator windings 210a-210c. The second depletion-mode transistor ensures that the braking resistor 206 can remain disconnected if one of the depletion-mode transistors (207 or 401) fails. Thus, power consumption and damage to motor system components can be avoided.

[0067] Figure 5 illustrates a motor control system 500 including a relay 501, consistent with embodiments of the present disclosure. Figure 5 includes the same devices as the control systems 200 in Figures 2 and 3, except that the switching device is a relay 501 instead of a transistor 207. In some embodiments, the relay 501 may be a single-pole double-throw (SDPT) relay, while in other embodiments, it may include any relay that enables switching between two circuits. The relay 501 may include an electromagnetic coil 501D, a common terminal 501C, a normally closed terminal 501A, and a normally open terminal 501B. Under load, the controller 201 may allow current to pass through the electromagnetic coil 501D, and the common terminal 501C may be connected to the normally open terminal 501A. Thus, under load, the braking resistor 206 remains disconnected. Under no-load conditions, the controller 201 may not allow current to pass through the electromagnetic coil 501D, and the common terminal 501C may be connected to the normally closed terminal 501B. Therefore, in the no-load state, the damping resistor 206 may be connected, allowing the generated current to pass across the terminals (501C and 501B) and through the damping resistor 206. Although only a single relay 501 is shown in Figure 5, the disclosure is not limited in that way. A second relay, similar to transistors 401 and 207 in Figure 4, may be included and provide the same redundancy benefits described above.

[0068] Figure 6 illustrates a motor control system 600 that includes a set of transistors 601a–601c. Figure 6 includes many of the same devices as the control system 200 in Figures 2 and 3 (e.g., power supply 204, capacitor 205, controller 201, sensor 203, and motor bridge 202). Similar to transistors 209a–c, transistors 601a–601c connect and disconnect the flow of current across the circuit to the stator windings 210a–210c based on whether voltage is received from the controller 201 and / or the motor bridge 202. In some embodiments, transistors 601a–601c are bipolar transistors, field-effect transistors (e.g., MOSFETs), or insulated-gate bipolar transistors (e.g., IGBTs). In some embodiments, transistors 601a-601c are depletion-mode transistors, and when the controller 201 and / or motor bridge 202 do not generate voltage across transistors 601a-601c, the transistors may allow current to flow across their terminals. In some embodiments, transistors 601a-601c are depletion-mode transistors, and transistors 208a-208c are enhancement-mode transistors. Thus, while the power supply 204 is connected under load, the controller 201 and / or motor bridge 202 may provide voltage across enhancement-mode transistors 208a-208c to allow current to flow, but may not provide voltage across depletion-mode transistors 601a-601c to allow current to flow. For example, in one embodiment, a voltage may be supplied to the enhancement mode transistor 208c to magnetize the stator windings 210c and 210a, but no voltage may be supplied to the depletion mode transistor 601a, and current may flow through transistor 208c, through stator winding 210c, through stator winding 210a, and through transistor 601a to form a completed circuit.

[0069] During the no-load state when the power supply 204 is disconnected, the controller 201 and / or motor bridge 202 may not provide voltage across enhancement-mode transistors 208a-208c, nor across depletion-mode transistors 601a-601c. Therefore, during the no-load state, current may flow through transistors 601a-601c rather than 208a-208c, as indicated by current line 212. This can result in a short circuit of all three motor phases, creating a counter-torque that opposes the rotation of the magnetic rotor (not shown) and connected flight elements.

[0070] Figure 7a illustrates a motor control system 700, consistent with embodiments of the present disclosure, which includes a diode 701, an enhancement-mode transistor 702, a passive discharge resistor 703, and an enhancement-mode transistor 704. As will be further detailed below, the switching device further comprises an enhancement-mode transistor 702 in addition to a depletion-mode transistor 207 to connect and disconnect the flow through the damping resistor 206. Figure 7a includes many of the same devices as the control system 200 in Figures 2 and 3 (e.g., power supply 204, damping resistor 206, capacitor 205, controller 201, sensor 203, and motor bridge 202). In some embodiments, transistors 702 and / or 704 are bipolar transistors, field-effect transistors (e.g., MOSFETs), and / or insulated-gate bipolar transistors (e.g., IGBTs). The enhancement-mode transistor 702 may, upon receiving a voltage from the controller 201, allow current to propagate across its terminals. The enhancement mode transistor 704 may allow current to flow across its terminals when it receives a voltage from the circuit through the passive discharge resistor 703. In some embodiments, the diode 701 may allow current to flow in both forward and reverse directions. The diode 701 may be doped to allow current to flow in the reverse direction only when a certain specified threshold breakdown voltage is reached. In some embodiments, the threshold voltage may be between 1.2 volts and 200 volts. When the threshold voltage is reached, the diode 701 may allow current to flow through its terminals. In some embodiments, the diode 701 may be a Zener diode, while in other embodiments, the diode 701 may be any device or circuit that allows current to flow in the reverse direction after the threshold voltage is reached. The passive discharge resistor 703 may be connected in series with the diode 701 to regulate the current to the diode 701. Under normal load conditions with the power supply 204 connected, the motor control system may drive the motor as described above with reference to Figure 2.

[0071] As shown in Figure 7A, under no-load conditions, the control system 700 may allow the energy stored in the bulk capacitor 205 to be discharged before turning off the flight control system and associated controller 201 and / or motor bridge 202. For example, the control system 700 may allow the stored energy to be discharged after flight for the safety of aircraft maintenance personnel. The controller 201 may supply voltage to the enhancement mode transistor 702, allowing current from the bulk capacitor 205 to flow through the passive discharge resistor 703 to the enhancement mode transistor 702, bypassing the Zener diode 701 and the damping resistor 206, and enabling the discharge of the bulk capacitor 205.

[0072] Figure 7b illustrates a motor control system including a passive discharge resistor and diode in a damped state, consistent with an embodiment of the present disclosure. As shown in Figure 7b, in a no-load state, after the power to the flight control system and associated controller 201 and / or motor bridge 202 is turned off, the control system 700 may dampen the motion of the flight elements. During the no-load state with the power supply 204 disconnected, stronger winds will increase the motion of the flight elements and the connected rotating magnetic rotor (not shown), and as the rotation increases, a larger inverse EMF (electromotive force) voltage may be induced in the circuit. The current generated from the inverse EMF voltage may be used to recharge the capacitor 205. As the inverse EMF voltage increases, the voltage across the Zener diode 701 may also increase. When the voltage across the Zener diode 701 exceeds the breakdown threshold voltage, current may flow across the passive discharge resistor 703, creating a voltage for the enhancement mode transistor 704. Next, the enhancement mode transistor 704 can allow the generated current to be transmitted across its terminals and through the damping resistor 206. As described above, the circuit connected through the stator windings 210a-210c and the damping resistor 206 provides a reverse torque that dampens the motion of the rotor (not shown) and connected flight elements. Furthermore, the damping resistor 206 allows excess energy to be dissipated in the form of heat.

[0073] The above description is provided for illustrative purposes only. It is not exhaustive and does not limit this disclosure to the exact form or embodiments disclosed. Modifications and adaptations of this disclosure will be apparent to those skilled in the art from the examination of this specification and the practice of the disclosed embodiments.

[0074] Figure 8 illustrates a method for mitigating short-circuit failures consistent with embodiments of the present disclosure. In some embodiments, the controller 201 performs the process shown in Figure 8, while in other embodiments, one or more different devices perform one or more steps of the process. In step 801, the controller 201 may detect a short-circuit condition through one or more voltage-sensing and / or current-sensing components. Furthermore, the voltage-sensing and / or current-sensing components may be arranged in parallel and / or series with one or more transistors to detect problems with each of the one or more transistors. For example, in some embodiments, the controller 201 may detect a short circuit associated with one or more transistors (208a-208c, 209a-209c). In some embodiments, the controller 201 and / or the battery management system may disconnect the power supply 204 (for example, by opening a switch and / or blowing a fuse). In step 803, after detecting a short circuit, the controller 201 and / or motor bridge 202 may instruct the transistors (208a-208c, 209a-209c) to implement a three-phase short circuit. For example, the short circuit may be associated with at least one of the transistors 209a-209b, and the controller 201 may instruct transistors 208a-208c to apply a three-phase short circuit. Similarly, the short circuit may be associated with at least one of the transistors 208a-208b, and the controller 201 may instruct transistors 209a-209c to apply a three-phase short circuit.

[0075] In step 807, the bus capacitor 205 is discharged. For example, in some embodiments, current may flow across the braking resistor 206 and discharge the bulk capacitor 205 using any one of the means that enable current to flow across the braking resistor 206, as previously described with reference to Figures 3-5 and 7A-7B. In some embodiments, the bulk capacitor 205 may be discharged through a passive discharge resistor separate from the braking resistor 206, as described above with reference to Figures 7A-7B. After the bulk capacitor 205 has started to discharge, the controller 201 may determine that the voltage across the bulk capacitor 205 has dropped below a threshold. In some embodiments, the controller 201 may receive information about the voltage across the bulk capacitor 205 from a different device (e.g., a battery management system) to make this determination. In some embodiments, the controller 201 may directly monitor the voltage across the bulk capacitor 205 using one or more voltage sensing devices and / or circuits. For example, the controller 201 may monitor the bus voltage across the capacitor 205 using a voltage comparator. In some embodiments, the motor circuit may be an inverter circuit (e.g., for a propeller, prop rotor, or rotor), and the voltage threshold may be about 40V to 60V. In some embodiments, the motor circuit may be a DC motor circuit (e.g., for a flight control plane), and the voltage threshold may be lower.

[0076] In step 809, based on the determination that the voltage across the bulk capacitor 205 has dropped below a threshold, the controller 201 may apply a counter-torque to counteract the motion of the flight element. The controller 201 may provide a counter-torque to counteract the motion of the flight element using any of the means described above with reference to Figures 3-6 and 7B. For example, in some embodiments, the controller 201 completely short-circuits the motor, as described with reference to Figure 6. In some embodiments, the controller 201 allows current to flow through the damping resistor 206, as described with reference to Figures 3-5 and 7B. For example, the controller 201 may close all transistors (e.g., 208a-208c and 209a-209c) and connect the damping resistor 206 circuit to allow current to flow through the damping resistor 206.

[0077] The above description is provided for illustrative purposes only. It is not exhaustive and does not limit the invention to the exact form or embodiment disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art from the examination herein and the practice of the embodiments disclosed herein.

[0078] The features and advantages of this disclosure are evident from the detailed specification, and therefore the attached claims are intended to cover all systems and methods that fall within the true intent and scope of this disclosure. As used herein, the indefinite articles "a" and "an" mean "one or more." Similarly, the use of plural terms does not necessarily indicate plural unless it is clear in the given context. Words such as "and" or "or" mean "and / or" unless otherwise indicated. Furthermore, since numerous modifications and variations can easily arise from examining this disclosure, it is not desirable to limit this disclosure to the exact configurations and operations illustrated and described, and therefore all suitable modifications and equivalents that fall within the scope of this disclosure may be applicable.

[0079] Other embodiments will be apparent to those skilled in the art from the examination of this specification and the practice of the implementations disclosed herein. The architectures and circuit layouts shown in the figures are intended for illustrative purposes only and are not intended to limit the invention to the specific configurations and circuit layouts shown in the figures. Furthermore, this specification and the examples are intended to be considered merely illustrative, and the true scope and spirit of the invention are set forth by the following claims. The above description is presented for illustrative purposes only. It is not exhaustive and does not limit the invention to the exact forms or embodiments disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art from the examination of this specification and the practice of the disclosed embodiments of the invention disclosed herein.

[0080] Embodiments of this disclosure may be further described by the following clauses. Clause 1. An electrical system for an aircraft, comprising a controller, a resistor, and a first transistor connected in series with the resistor, wherein the controller is configured to control the first transistor to prevent the flow of current across the terminals of the first transistor and through the resistor when the aircraft's power supply is providing power, and the first transistor is configured to allow a current generated by a reverse EMF (electromotive force) voltage to flow across the terminals of the first transistor and through the resistor when the aircraft's power supply is not providing power, the reverse EMF voltage being generated by the movement of the aircraft's movable aircraft structure. 2. An electrical system as described in Clause 1, further comprising a resistor and a second transistor connected in series with the first transistor. 3. An electrical system as described in Clause 2, wherein the controller is configured to control the second transistor to prevent the flow of current across the terminals of the second transistor and through the resistor when the power supply of the aircraft is providing power, and the second transistor is configured to allow the current generated by the reverse EMF voltage to flow across the terminals of the second transistor and through the resistor when the power supply of the aircraft is not providing power. 4. An electrical system as described in any one of Clauses 1 to 3, wherein the controller is configured to control the first transistor by supplying a voltage to the first transistor to prevent the flow of current across the terminals of the first transistor and through the resistor. 5. An electrical system as described in any one of clauses 1 to 4, wherein the movable aircraft structure is a propeller or a flight control surface. 6. An electrical system as described in any one of Clauses 1 to 5, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor, each of the third, fourth, fifth, and sixth transistors configured to control the flow of current from the power supply of the aircraft to the stator windings of the aircraft's motors in order to control the movable aircraft structure. 7. An electrical system as described in Clause 6, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor is configured to allow current to flow across the terminals of the at least one transistor and through the resistor when the power supply of the aircraft is not providing power. 8. An electrical system as described in Clause 7, wherein the at least one transistor is configured to allow the flow of current across the terminals of the at least one transistor and through the resistor when the at least one transistor is not receiving a voltage. 9. An electrical system according to any of clauses 6 to 8, wherein the movable aircraft structure comprises a propeller, and the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor are configured to provide a sinusoidal voltage output to the stator winding of the motor of the aircraft in order to control the propeller. 10. An electrical system as described in any one of clauses 6 to 9, further comprising a capacitor, wherein the controller is configured to detect a short circuit and control the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor, the control of which the third, fourth, fifth, and sixth transistors short-circuits all of the stator windings, and the capacitor discharges through the resistor when the controller short-circuits the stator windings. 11. An electrical system as described in Clause 10, wherein the controller is configured to detect when the capacitor has discharged below a threshold, and when the controller detects that the capacitor has discharged below the threshold, it is configured to allow the current generated by the reverse EMF (electromotive force) voltage to flow through the resistor. 12. An aircraft equipped with an electrical system as described in any of clauses 1 to 11. 13. An aircraft as described in Article 12, wherein the aircraft is an electric aircraft. 14. An electrical system for an aircraft, comprising a controller, a resistor, and a relay connected in series with the resistor, wherein the controller is configured to control the relay to prevent the flow of current across the resistor when the aircraft's power supply is providing power, and the relay is configured to allow a current generated by a reverse EMF (electromotive force) voltage to flow across the resistor when the aircraft's power supply is not providing power, the reverse EMF voltage being generated by the movement of the aircraft's movable aircraft structure. 15. An electrical system as described in Article 14, wherein the relay is a single-pole double-throw relay. 16. An electrical system as described in Clause 14 or 15, wherein the controller is configured to control the relay to prevent the flow of current through the resistor by supplying voltage to the relay. 17. An electrical system as described in any one of clauses 14 to 16, wherein the movable aircraft structure is a propeller or a flight control surface. 18. An electrical system as described in any one of Clauses 14 to 17, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor configured to control the flow of current to the stator windings of the aircraft's motor for controlling the movable aircraft structure. 19. An electrical system as described in Clause 18, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor is configured to allow current to flow across the terminals of the at least one transistor and through the resistor when the power supply of the aircraft is not providing power. 20. An electrical system as described in Clause 19, wherein the at least one transistor is configured to allow the flow of current across the terminals of the at least one transistor and through the resistor when the at least one transistor is not receiving a voltage. 21. An electrical system according to any of the clauses 18 to 20, wherein the movable aircraft structure comprises a propeller, and the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor are configured to provide a sinusoidal voltage output to the stator winding of the motor of the aircraft in order to control the propeller. 22. An electrical system as described in any one of clauses 18 to 21, further comprising a capacitor, wherein the controller is configured to detect a short circuit and control the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor to short-circuit all of the stator windings, and the capacitor discharges through the resistor when the controller short-circuits the stator windings. 23. An electrical system as described in Clause 22, wherein the controller is configured to detect when the capacitor has discharged below a threshold, and when the controller detects that the capacitor has discharged below the threshold, it is configured to allow the current generated by the reverse EMF (electromotive force) voltage to flow through the resistor. 24. An aircraft equipped with an electrical system as described in any of clauses 14 to 23. 25. An aircraft as described in Article 25, wherein the aircraft is an electric aircraft. 26. An electrical system for an aircraft, comprising a controller, a first set of transistors, and a second set of transistors, wherein the controller is configured to control the first set of transistors and the second set of transistors to create a magnetic field in the stator windings of the aircraft's motors when the aircraft's power supply is providing power, the first set of transistors is configured to prevent current generated by an inverse EMF (electromotive force) voltage from flowing across the terminals of the first set of transistors when the aircraft's power supply is not providing power, the inverse EMF voltage being generated by the movement of a movable aircraft structure, and the second set of transistors is configured to allow current generated by an inverse EMF voltage to flow across the terminals of the second set of transistors when the aircraft's power supply is not providing power. 27. An electrical system as described in Clause 26, wherein the second set of transistors is a depletion-mode transistor. 28. An electrical system as described in Clause 26 or 27, wherein the first set of transistors are enhancement-mode transistors. 29. An electrical system as described in any one of clauses 26 to 28, wherein the first set of transistors is connected in series with the second set of transistors. 30. An electrical system as described in any one of clauses 26 to 29, wherein the aircraft structure comprises a propeller or a flight control surface. 31. An electrical system as described in any one of Clauses 26 to 30, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor configured to control the flow of current to the stator windings for controlling the movable aircraft structure. 32. An electrical system as described in Clause 31, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor is configured to allow current to flow across the terminals of the at least one transistor and across a resistor when the power supply of the aircraft is not providing power. 33. An electrical system as described in Clause 32, wherein the at least one transistor is configured to allow current to flow across the terminals of the at least one transistor and across a resistor when the at least one transistor is not receiving a voltage. 34. An electrical system according to any of the clauses 31 to 33, wherein the movable aircraft structure comprises a propeller, and the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor are configured to provide a sinusoidal voltage output to the stator winding of the aircraft's motor for controlling the propeller. 35. An electrical system as described in any one of clauses 31 to 34, further comprising a capacitor, wherein the controller is configured to detect a short circuit and control the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor to short-circuit all of the stator windings, and the capacitor discharges through the resistor when the controller short-circuits the stator windings. 36. An electrical system as described in Clause 35, wherein the controller is configured to detect when the capacitor has discharged below a threshold, and when the controller detects that the capacitor has discharged below the threshold, it is configured to allow the current generated by the reverse EMF (electromotive force) voltage to flow through the resistor. 37. An aircraft equipped with an electrical system as described in any of clauses 26 to 36. 38. An aircraft as described in Article 37, wherein the aircraft is an electric aircraft. 39. An electrical system for an aircraft, comprising a controller, a first resistor, and a first transistor connected in series with the first resistor, wherein the first transistor is configured to prevent the flow of current across its terminals and through the first resistor when the aircraft's power supply is providing power, and the first transistor is configured to allow a current generated by a reverse EMF (electromotive force) voltage to flow across its terminals and through the first resistor when the power supply is not providing power, the reverse EMF voltage being generated by the movement of the aircraft's movable aircraft structure. 40. An electrical system as described in Article 39, further comprising a diode and a second resistor. 41. An electrical system as described in Clause 40, wherein the diode includes a threshold voltage, and when the reverse EMF voltage exceeds the threshold voltage, the diode allows current to flow through the second resistor. 42. An electrical system as described in Clause 41, wherein the flow of current through the diode and the second resistor enables the first transistor to carry out a current generated by the reverse EMF voltage across the terminals of the first transistor and through the first resistor. 43. An electrical system as described in any one of clauses 39 to 42, further comprising a second transistor, wherein the controller is configured to control the second transistor so as to allow current to flow through the terminals of the second transistor and to allow a capacitor to discharge. 44. An electrical system as described in any one of clauses 39 to 43, wherein the reverse EMF voltage is produced by the movement of the propeller or flight control surface. 45. An electrical system as described in any one of clauses 39 to 44, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor, each of the third, fourth, fifth, and sixth transistors configured to control the flow of current from the power supply of the aircraft to the stator windings in order to control the movable aircraft structure. 46. ​​An electrical system as described in Clause 45, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor is configured to allow current to flow through the terminals of the at least one transistor and the first resistor when the power supply of the aircraft is not providing power. 47. An electrical system as described in Clause 46, wherein the at least one transistor is configured to allow the flow of the current across the terminals of the at least one transistor and through the first resistor when the at least one transistor is not receiving a voltage. 48. An electrical system as described in any one of clauses 45 to 47, wherein the movable aircraft structure comprises a propeller, and the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor are configured to provide a sinusoidal voltage output to the stator winding of the motor for controlling the propeller. 49. An electrical system as described in any one of clauses 45 to 48, further comprising a capacitor, wherein the controller is configured to detect a short circuit and to control the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor, each of the third, fourth, fifth, and sixth transistors being configured to control the flow of current from the power supply of the aircraft to the stator windings in order to control the movable aircraft structure. 50. An electrical system as described in Clause 49, wherein the controller is configured to detect when the capacitor has discharged below a threshold, and the controller is configured to allow the current generated by the reverse EMF (electromotive force) voltage to flow through the first resistor when it detects that the capacitor has discharged below the threshold. 51. A method for damping the motion of a movable aircraft structure, comprising controlling a switching device by a controller to prevent the flow of current through a first resistor when the aircraft's power supply is providing power, wherein the switching device allows a current generated by a reverse EMF (electromotive force) voltage to flow through the first resistor when the aircraft's power supply is not providing power, the reverse EMF voltage being generated by the motion of the movable aircraft structure of the aircraft. 52. A method according to Article 51, wherein the switching device is a first transistor. 53. The method according to Clause 52, wherein the second transistor is connected in series with the first resistor and the first transistor. 54. A method according to Clause 53, further comprising controlling the second transistor by the controller to prevent the flow of current across the terminals of the second transistor and through the first resistor when the power supply of the aircraft is providing power, wherein the second transistor allows current generated by the reverse EMF voltage to flow across the terminals of the second transistor and through the first resistor when the power supply of the aircraft is not providing power. 55. A method according to Clause 52 or 53, wherein the controller controls the first transistor by providing a voltage to the first transistor to prevent the flow of current across the terminals of the first transistor and through the first resistor. 56. A method according to any of the provisions 51 to 55, wherein the inverse EMF voltage is produced by the movement of the propeller or flight control surface. 57. A method according to any of the provisions 51 to 56, further comprising controlling the flow of current from the power supply of the aircraft to the stator windings of the aircraft's motors in order to control the movable aircraft structure by a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor. 58. A method according to Clause 57, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor allows current to flow across the terminals of the at least one transistor and through the first resistor when the power supply of the aircraft is not providing power. 59. An electrical system as described in Clause 58, wherein the at least one transistor enables the flow of current across the terminals of the at least one transistor and through the first resistor when the at least one transistor does not receive a voltage from the controller. 60. A method according to any of the provisions 57 to 59, wherein the movable aircraft structure comprises a propeller, and the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor provide a sinusoidal voltage output to the stator winding of the motor to control the propeller. 61. A method according to any of the provisions 57 to 60, wherein the controller detects a short circuit and controls the third, fourth, fifth, and sixth transistors, thereby short-circuiting all of the stator windings, and the capacitor discharges through the first resistor when the controller short-circuits the stator windings. 62. A method according to Clause 61, wherein the controller detects when the capacitor has discharged below a threshold, and when the controller detects that the capacitor has discharged below the threshold, it allows the current generated by the reverse EMF (electromotive force) voltage to flow through the first resistor. 63. A method according to Article 51, wherein the switching device is a relay. 64. A method according to Article 63, wherein the relay is a single-pole double-throw relay. 65. A method according to Clause 63 or 64, wherein the controller controls the relay to prevent the flow of current through the first resistor by providing a voltage to the relay. 66. A method according to any of the provisions 63 to 65, wherein the inverse EMF voltage is produced by the movement of the propeller or flight control surface. 67. A method according to any of the provisions 63 to 66, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor, each of the third, fourth, fifth, and sixth transistors controlling the flow of current to the stator windings of the aircraft's motor to control the movable aircraft structure. 68. A method according to Clause 67, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor allows current to flow across the terminals of the at least one transistor and through the first resistor when the power supply of the aircraft is not providing power. 69. An electrical system as described in Clause 68, wherein the at least one transistor enables the flow of current across the terminals of the at least one transistor and through the first resistor when the at least one transistor does not receive a voltage from the controller. 70. A method according to any of the provisions 67 to 69, wherein the movable aircraft structure comprises a propeller, and the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor provide a sinusoidal voltage output to the stator winding of the motor in order to control the propeller. 71. A method according to any of the provisions 67 to 70, wherein the controller detects a short circuit and controls the third, fourth, fifth, and sixth transistors, thereby short-circuiting all of the stator windings, and the capacitor discharges through the first resistor when the controller short-circuits the stator windings. 72. A method according to Clause 61, wherein the controller detects when the capacitor has discharged below a threshold, and when the controller detects that the capacitor has discharged below the threshold, it allows the current generated by the reverse EMF (electromotive force) voltage to flow through the first resistor. 73. The method according to Article 52, further comprising a diode and a second resistor. 74. A method according to Clause 73, wherein the reverse EMF voltage causes the diode to exceed a breakdown voltage threshold, and the diode allows current to flow through the second resistor. 75. A method according to Clause 74, wherein the flow of current through the diode and the second resistor allows the first transistor to allow the current generated by the reverse EMF voltage to flow through the first resistor. 76. A method according to any of the provisions 73 to 75, wherein the controller controls the second transistor to allow current to flow through the terminals of the second transistor and to allow the capacitor to discharge. 77. A method according to any of the provisions 73 to 76, comprising controlling the flow of current from the power supply of the aircraft to the stator windings in order to control the movable aircraft structure, by each of the third transistor, fourth transistor, fifth transistor, and sixth transistor. 78. A method according to Article 77, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor allows current to flow across the terminals of the at least one transistor and through the first resistor when the power supply of the aircraft is not providing power. 79. A method according to Clause 77 or 78, wherein the movable aircraft structure comprises a propeller, and the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor provide a sinusoidal voltage output to a stator winding to control the propeller. 80. A method according to any of the provisions 77-79, wherein the controller detects a short circuit and controls the third, fourth, fifth, and sixth transistors, thereby short-circuiting all of the stator windings, and the capacitor discharges through the second resistor when the controller short-circuits the stator windings, and the controller detects when the capacitor has discharged below a threshold, and when the controller detects that the capacitor has discharged below the threshold, the controller allows the current generated by the reverse EMF (electromotive force) voltage to flow through the first resistor. 81. An electrical system configured to implement any of the methods described in clauses 51 to 80. 82. An aircraft equipped with an electrical system as described in any of clauses 39 to 50. 83. An aircraft as described in Article 82, wherein the aircraft is an electric aircraft.

[0081] The above description is provided for illustrative purposes only. It is not exhaustive and does not limit the invention to the exact form or embodiment disclosed herein. Modifications and adaptations of the invention will be apparent to those skilled in the art from the examination herein and the practice of the disclosed embodiments of the invention disclosed herein.

Claims

1. An electrical system for aircraft, controller, The first resistor, The first resistor is connected in series with a switching device, The controller is configured to control the switching device to prevent the flow of current through the first resistor when the aircraft's power supply is providing power, The switching device is configured to allow a current generated by a reverse EMF (electromotive force) voltage to flow through the first resistor when the power supply of the aircraft is not providing power, and the reverse EMF voltage is generated by the movement of the aircraft's movable aircraft structure, in an electrical system.

2. An electrical system according to claim 1, wherein the switching device is a first transistor.

3. An electrical system according to claim 2, further comprising a second transistor connected in series with the first resistor and the first transistor.

4. The electrical system according to claim 3, The controller is configured to control the second transistor to prevent the flow of current across the terminals of the second transistor and through the first resistor when the power supply of the aircraft is providing power. An electrical system in which the second transistor is configured to allow a current generated by the inverse EMF voltage to flow across the terminals of the second transistor and through the first resistor when the power supply of the aircraft is not providing power.

5. An electrical system according to claim 2 or 3, wherein the controller is configured to control the first transistor by supplying a voltage to the first transistor to prevent the flow of current across the terminals of the first transistor and through the first resistor.

6. An electrical system according to any one of claims 1 to 5, wherein the movable aircraft structure is a propeller or a flight control surface.

7. An electrical system according to any one of claims 1 to 6, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor, each of the third, fourth, fifth, and sixth transistors configured to control the flow of current from the power supply of the aircraft to the stator windings of the aircraft's motor in order to control the movable aircraft structure.

8. An electrical system according to claim 7, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor is configured to allow current to flow across the terminals of the at least one transistor and through the first resistor when the power supply of the aircraft is not providing power.

9. An electrical system according to claim 8, wherein the at least one transistor is configured to allow the flow of current across the terminals of the at least one transistor and through the first resistor when the at least one transistor does not receive a voltage from the controller.

10. An electrical system according to any one of claims 7 to 9, The aforementioned movable aircraft structure is equipped with a propeller, An electrical system in which the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor are configured to provide a sinusoidal voltage output to the stator winding of the motor in order to control the propeller.

11. An electrical system according to any one of claims 7 to 10, Equipped with additional capacitors, The controller is configured to detect a short circuit and control the third, fourth, fifth, and sixth transistors, and controlling the third, fourth, fifth, and sixth transistors short-circuits all of the stator windings. An electrical system in which the capacitor discharges through the first resistor when the controller short-circuits the stator winding.

12. The electrical system according to claim 11, The controller is configured to detect when the capacitor discharges below a threshold, An electrical system in which, when the controller detects that the capacitor has discharged below the threshold, it allows the current generated by the reverse EMF (electromotive force) voltage to flow through the first resistor.

13. An electrical system according to claim 1, wherein the switching device is a relay.

14. An electrical system according to claim 13, wherein the relay is a single-pole double-throw relay.

15. An electrical system according to claim 13 or 14, wherein the controller is configured to control the relay so as to prevent the flow of current through the first resistor by supplying a voltage to the relay.

16. An electrical system according to any one of claims 13 to 15, wherein the movable aircraft structure is a propeller or a flight control surface.

17. An electrical system according to any one of claims 13 to 16, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor, each of the third, fourth, fifth, and sixth transistors configured to control the flow of current to the stator windings of the aircraft's motor in order to control the movable aircraft structure.

18. An electrical system according to claim 17, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor is configured to allow current to flow across the terminals of the at least one transistor and through the first resistor when the power supply of the aircraft is not providing power.

19. An electrical system according to claim 18, wherein the at least one transistor is configured to allow the flow of current across the terminals of the at least one transistor and through the first resistor when the at least one transistor does not receive a voltage from the controller.

20. An electrical system according to any one of claims 17 to 19, The aforementioned movable aircraft structure is equipped with a propeller, An electrical system in which the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor are configured to provide a sinusoidal voltage output to the stator winding of the motor in order to control the propeller.

21. An electrical system according to any one of claims 17 to 20, Equipped with additional capacitors, The controller is configured to detect a short circuit and control the third, fourth, fifth, and sixth transistors, and controlling the third, fourth, fifth, and sixth transistors short-circuits all of the stator windings. An electrical system in which the capacitor discharges through the first resistor when the controller short-circuits the stator winding.

22. The electrical system according to claim 21, The controller is configured to detect when the capacitor discharges below a threshold, An electrical system in which, when the controller detects that the capacitor has discharged below the threshold, it allows the current generated by the reverse EMF (electromotive force) voltage to flow through the first resistor.

23. An electrical system according to claim 2, further comprising a diode and a second resistor.

24. An electrical system according to claim 23, wherein the diode includes a threshold voltage, and when the reverse EMF voltage exceeds the threshold voltage, the diode allows current to flow through the second resistor.

25. An electrical system according to claim 24, wherein the flow of current through the diode and the second resistor enables the first transistor to allow the current generated by the reverse EMF voltage to flow through the first resistor.

26. An electrical system according to any one of claims 23 to 25, further comprising a second transistor, wherein the controller is configured to control the second transistor so as to allow current to flow through the terminals of the second transistor and to allow a capacitor to discharge.

27. An electrical system according to any one of claims 23 to 26, further comprising a third transistor, a fourth transistor, a fifth transistor, and a sixth transistor, each of the third, fourth, fifth, and sixth transistors configured to control the flow of current from the power supply of the aircraft to the stator windings in order to control the movable aircraft structure.

28. An electrical system according to claim 27, wherein at least one of the third transistor, the fourth transistor, the fifth transistor, or the sixth transistor is configured to allow current to flow across the terminals of the at least one transistor and through the first resistor when the power supply of the aircraft is not providing power.

29. An electrical system according to claim 27 or 28, The aforementioned movable aircraft structure is equipped with a propeller, An electrical system in which the third transistor, the fourth transistor, the fifth transistor, and the sixth transistor are configured to provide a sinusoidal voltage output to the stator windings in order to control the propeller.

30. An electrical system according to any one of claims 27 to 29, Equipped with additional capacitors, The controller is configured to detect a short circuit and control the third, fourth, fifth, and sixth transistors, and controlling the third, fourth, fifth, and sixth transistors short-circuits all of the stator windings. When the controller short-circuits the stator winding, the capacitor discharges through the second resistor. The controller is configured to detect when the capacitor discharges below a threshold, An electrical system in which, when the controller detects that the capacitor has discharged below the threshold, it allows the current generated by the reverse EMF (electromotive force) voltage to flow through the first resistor.

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