WO2024127325A1

WO2024127325A1 - Motor controller, control system, and a method for operating the control system

Info

Publication number: WO2024127325A1
Application number: PCT/IB2023/062722
Authority: WO
Inventors: Sébastien DEMONT; Milan UTVIC; David COSTES; Aurélien CARRUPT
Original assignee: H55 Sa
Priority date: 2022-12-15
Filing date: 2023-12-14
Publication date: 2024-06-20
Also published as: WO2024127319A1; WO2024127324A1; WO2024127322A1; WO2024127323A1

Abstract

A motor controller (93) for an electrically or hybrid driven aircraft (100), comprising: - an input end configured to receive electrical energy from an electrical source (91, 92); - an output end configured to supply electrical energy to and receive electrical energy from a motor (94); - a signal input configured to receive signals from a speed sensor (960), wherein the electrical source (91, 92), the motor (94), and the speed sensor (960) are external to the motor controller (93), and wherein the motor controller (93) is arranged to provide electrical energy at the output end in the form of a set of driving signals with variable voltage and variable frequency with the use of different control modes for controlling the set of driving signals, wherein in a sensor-based control mode the motor controller (93) is configured to control the set of driving signals at the output end based on a first control scheme, wherein the first control scheme is configured to utilize the signals at the signal input and in a sensorless control mode the motor controller (93) is configured to control the set of driving signals with the use of a second control scheme where the signals at the signal input are disregarded.

Description

Motor controller, control system, and a method for operating the control system Technical domain [0001] The present disclosure concerns a motor controller and a control system for an electrically or hybrid driven aircraft. The disclosure also concerns a method for controlling the control system and an electrically or hybrid driven aircraft with the control system controlled by the method. [0002] The motor controller and the control system are arranged for maintaining the operation of the motor controller and the aircraft propelled by the control system, even in case of a failure of a speed sensor. Related art [0003] Electric and hybrid vehicles have become increasingly significant for the transportation of people and goods. Such vehicles can desirably provide energy efficiency advantages over combustion-powered vehicles and may cause less air pollution than combustion-powered vehicles during operation. [0004] Although the technology for electric and hybrid automobiles has significantly developed in recent years, many of the innovations that enabled a transition from combustion-powered to electric-powered automobiles unfortunately do not directly apply to the development of electric or hybrid aircraft. The functionality of automobiles and the functionality of aircraft are sufficiently different in many aspects so that many of the design elements for electric and hybrid aircraft must be uniquely developed separate from those of electric and hybrid automobiles.

H55-24-PCT [0005] Moreover, any changes to an aircraft's design, such as to enable electric or hybrid operation, also require careful development and testing to ensure safety and reliability. If an aircraft experiences a serious failure during flight, the potential loss and safety risk from the failure may be very high as the failure could cause a crash of the aircraft and pose a safety or property damage risk to passengers or cargo, as well as individuals or property on the ground. [0006] The certification standards for electric or hybrid aircraft are further extremely stringent because of the risks posed by new aircraft designs. Designers of aircraft have struggled to find ways to meet the certification standards and bring new electric or hybrid aircraft designs to market. [0007] In view of these challenges, attempts to make electric and hybrid aircraft commercially viable have been largely unsuccessful. New approaches for making and operating electric and hybrid aircraft thus continue to be desired. [0008] Flying a manned or unmanned aircraft such an airplane can be dangerous. Problems with the aircraft may result in injury or loss of life for passengers in the aircraft or individuals on the ground, as well as damage to goods being transported by the aircraft or other items around the aircraft. [0009] The reliability of systems can be improved with redundant subsystems. Various designs have been suggested in order to replace a faulty subsystem with a backup subsystem. For example, in the context of electric powered object or vehicles, US20171210229 A1 and US20111254502A1 both describe a fault-tolerant battery management system in which the state of battery cells is monitored and/or controlled by redundant battery management systems (BMS), such that a default in one BMS does not prevent the battery from functioning as long as the redundant BMS performs properly. However, if the two BMS are identical,

H55-24-PCT they are more likely to present the same defaults or conception problems, and are also more likely to have failure simultaneously or at short interval. Moreover, those solutions have not been designed with the aim of certification for aircraft; adding additional components increase the complexity of the system and makes the certification even more difficult. [0010] In order to attempt to mitigate potential problems associated with an aircraft, numerous organizations have developed certification standards for ensuring that aircraft designs and operations satisfy threshold safety requirements. The certification standards may be stringent and onerous when the degree of safety risk is high, and the certification standards may be easier and more flexible when the degree of safety risk is low. [0011] As an example, the FAA advisory circular AC 25.1309-1 describes acceptable means for showing compliance with the airworthiness requirements of US Federal Aviation Regulations defines different levels of failure conditions according to their severity: ^ Failure Conditions with No Safety Effect. ^ Minor Failure Conditions. ^ Major Failure Conditions. ^ Hazardous Failure Conditions must be no more frequent than Extremely Remote. ^ Catastrophic Failure Conditions must be Extremely Improbable. [0012] While airplanes must be designed so that hazardous and catastrophic failure conditions are extremely remote or even extremely improbable, those severe failure conditions must nevertheless be monitored, so that warning signals are sent to the pilot and driver who may attempt to remedy to the condition or try to land the aircraft. The monitoring and warning systems must be reliable and also requires certification.

H55-24-PCT [0013] Such certification standards have unfortunately had the effect of slowing commercial adoption and production of electric or hybrid aircraft. Electrical hybrid aircraft may, for example, utilize new aircraft designs relative to traditional aircraft designs to account for differences in operations of electric or hybrid aircraft versus traditional aircraft. The new designs however may be significantly different from the traditional aircraft designs. These differences may subject the new designs to extensive testing prior to certification. The need for extensive testing can take many resources, time and significantly drive up the ultimate cost of the aircraft. [0014] Compliance of a monitoring and warning subsystem with the certification standard depends on the severity of the monitored failure condition. Therefore, a hazardous or catastrophic failure condition requires a strict level of certification of the corresponding monitoring and warning system, while a minor failure condition or a condition without any safety effect have lower safety requirements and requires a monitoring and warning system that is easier to certify, or requires no certification. [0015] There is therefore a need for simplified, yet robust, components and systems for an electric powered aircraft that simplify and streamline certifications requirements and reduce the cost and time required to produce a commercially viable electric aircraft. [0016] Subsystems for propelling the electric aircraft are one of the most critical subsystems in an electric or hybrid aircraft, as a loss of propulsion might lead to catastrophic scenarios. Various solutions are known from prior art to prevent a loss of propulsion. Sensing circuits are important to identify failures in the subsystems. It is essential to identify a failure as early as possible and to react immediately with appropriate measures to mitigate the failure before it leads to a catastrophic scenario. [0017] Speed sensors, for instance, are crucial sensors of the propulsion subsystem. Conventionally they are used to determine the rotor position and the corresponding rotor speed. A failure of the speed sensor often

H55-24-PCT leads to the non-operability of the motor controller and the motor, and might lead to a loss of propulsion. Solutions known from the prior art often make use of a second redundant speed sensor. Suppose the main speed sensor fails, the motor controller can take benefit of the second speed sensor to continue the operation of the motor controller. However, the installation of two redundant speed sensors leads to additional costs, weight, and increased complexity of the propulsion system, which is in most cases not desirable. Moreover, the effort for certification will be increased, as it must be proven that the two redundant speed sensors do not interfere with each other during operation. [0018] The disclosure, therefore, has the objective of remedying some disadvantages of the prior art. In particular, it is an objective to provide a motor controller and a related control system for an electrically or hybrid driven aircraft that overcomes the disadvantaged. In particular, the motor controller and the related control system can continue the operation independently from a failure of a speed sensor. [0019] The disclosure significantly increases the safe operation and reduces the risk of catastrophic scenarios, which can be caused by a loss propulsion. The disclosed device is lightweight and cost-efficient, as no additional equipment, such as a second redundant sensor, is required. The objective is supplemented by a method for operating the control system. The method can be efficiently implemented in a control means of the motor controller. The simplicity of the method can significantly reduce the effort for certification. Short disclosure [0020] According to one aspect, a motor controller is disclosed, involving the features recited in claim 1. Further features and embodiments of the motor controller are described in the dependent claims.

H55-24-PCT [0021] The disclosure relates to a motor controller for an electrically or hybrid driven aircraft, comprising: - an input end configured to receive electrical energy from an electrical source; - an output end configured to supply electrical energy to and receive electrical energy from a motor; - a signal input configured to receive signals from a speed sensor, wherein the electrical source, the motor, and the speed sensor are external to the motor controller, and wherein the motor controller is arranged to provide electrical energy at the output end in the form of a set of driving signals with variable voltage and variable frequency with the use of different control modes for controlling the set of driving signals, wherein in a sensor-based control mode (called “the first control mode”) the motor controller is configured to control the set of driving signals at the output end based on a first control scheme, wherein the first control scheme is configured to utilize the signals at the signal input and in a sensorless control mode (the “second control mode”) the motor controller is configured to control the set of driving signals with the use of a second control scheme where the signals at the signal input are disregarded. [0022] The motor controller provides the advantage that it can provide the set of driving signals to the motor with the use of two independent control schemes, in which the first control scheme makes use of the speed sensor signals, and in which the second control scheme disregards the speed sensor signals. The functioning of the motor controller and the corresponding operation of the motor is ensured independently of the presence of a speed sensor. Suppose the speed sensor fails during the operation of the motor controller, i.e. during flight, the motor controller can advantageously unperturbed continue its operation. [0023] The electrical source can be any kind of electrical source capable of providing electric energy to the motor controller. For instance, batteries, or fuel electric generators can be used as an electrical source. The output end can be configured to supply electrical energy to or receive electrical

H55-24-PCT energy from the motor. However, the motor can be external to the motor controller, and thus can be connectable to the motor controller. The output end can also be referred to input-output end for better describe the functioning. [0024] The speed or position sensor can be any device that is suitable to provide the motor controller with the position or the speed (angular frequency) of the rotor of the motor. The position provided by the speed sensor can be relative or absolute. The speed or position sensor can be part of the motor controller or can preferably belong to the motor. In this case, the speed or position sensor can be considered as being external to the motor controller. However, in any case the motor controller can be configured to receive speed sensor signals. The signal input can be configurated as a dedicated interface to connect a speed or position sensor. The interface can be provided as a hard-wired connection to the speed or position sensor or as an interface connected to a communication bus. The term "receivable" with respect to the sensor signals at the signal input can imply that the sensors provide the related signal, in case the rotor or the motor rotates, at least in the absence of a defect. [0025] The set of driving signals can be considered as AC voltages and/or currents with variable frequency and variable amplitude. The frequency and amplitude of each driving signal can be varied depending on the demanded power (torque and speed) of the motor for propelling the airplane. The frequency and amplitude of the driving signals can be controlled by the motor controller using different control schemes. The set of driving signals preferably refers to a plurality of driving signals, such as AC voltages or currents in a three-phase system. Each driving signal can be mutually different in its phase but can be similar in frequency and amplitude. [0026] In an embodiment, the motor controller can be configured to determine or estimate a rotor position of the rotor by determining a stator current or voltage when the rotor rotates, and the motor generates

H55-24-PCT electrical energy and can be further arranged to initialize the sensorless control mode based on the determined or estimated rotor position. This embodiment can provide the advantage that the motor controller can initiate the sensorless control mode (second control mode) while the propeller is turning the rotor, due to windmill, or due to the rotor's inertia, etc. Initiate could mean start in sensorless control mode, or transition to the sensorless control mode. By determining or estimating the rotor position, the motor controller can seamlessly enter from the sensor-based into the sensorless control mode without risking overcurrent or voltage. If the rotor (the poles thereof) is in an unknown position relative to the stator windings, there is a risk of high initial current flow. Therefore, if the signal from the speed or position sensor is lost or erroneous, it can be useful to determine or estimate the rotor position in another way, particularly by determining the current or voltage outputted by the motor while the rotor is turning. [0027] In an embodiment , the second control scheme can be based on a V/f- or I/f-control, whereby V can denote the voltage, I can denote the current, and f can denote the frequency of the set of driving signals at the output end. V and/or I preferably denote the amplitude of the AC voltage and/or current. The V/f- or I/f-control can be used to start the motor in case the signal of the speed sensor is not available, i.e. in case of a failure. In particular, the V/f- or I/f-control can be used, as less torque can be required at start-up compared to conventional vehicles, such as cars or trucks. [0028] In another embodiment, the first control scheme can be based on sensor-based vector control, and the second control scheme can be based on sensorless vector control. In particular, the sensor-based vector control scheme can be used for utilizing the sensor signals. However, both control methods (sensorless and sensor-based) can result in a performant motor control, in which the speed and the corresponding torque of the motor can be precisely controlled. Sensorless relates to the fact that no speed or position sensor signal is used for determining the position or speed of the rotor, and sensor-based relates to the circumstance that a speed or position sensor signal is used.

H55-24-PCT [0029] The motor controller can also be configured with an initialization control mode (the “third control mode”), in which the motor controller can be arranged to receive electrical energy at the output end and can be configured to determine or estimate a rotor position from the electrical energy received. The motor can supply electrical energy to the motor controller, whereby the motor controller can be configured to determine the rotor position out of the electrical energy that is supplied by the motor. The electrical energy can be supplied in the form of signals that can be equivalent to the set of driving signals, in particular AC voltages and/or currents that are mutually different in their phase but similar in their amplitude and frequency. Initialization can mean that the motor controller does not provide electrical energy for propulsion but remains in the control mode for a short time to prepare the transition from the sensor-based control mode into the sensorless control mode, e.g. by synchronizing the set of driving signals with the rotor position determined or estimated. [0030] The virtual rotor position can preferably correspond to the rotor position of the motor as it would be determined with the use of the speed or position sensor and thus can correspond to the real rotor position of the motor. In this control mode, the motor controller can make use of the so- called windmill effect, in which the rotor remains rotating due to the airstream passing by the propeller and/or due to its inerty, and the motor thus continues to supply electrical energy in the form of said signals caused by the motors counter-electromotive force (back-EMF). [0031] In another embodiment, the motor controller can detect a failure at the signal input, wherein the motor controller can transition from the first control mode into the second control mode upon the detection of the failure. The failure can be caused for example by a defective sensor or by the loss of the connection of speed sensor at the signal input. The motor controller can also be configured to determine erroneous sensor signals. [0032] In a further embodiment, the motor controller can be configured to detect a failure at the signal input, wherein the motor controller can be

H55-24-PCT arranged to transition from the sensor-based, into the initialization and subsequently into the sensorless control mode upon the detection of the failure. [0033] In addition, the motor controller can be configured to utilize the rotor position determined or estimated for the second control scheme while switching or transitioning from the sensor-based or initialization control mode into the sensorless control mode. In particularly, this can be useful, as the determined or estimated rotor position can be used for initialization of the sensorless vector control. The rotor position, at least in the initializating phase, is required for the sensorless vector control to guarantee a smooth transition between the said control modes. Otherwise, the transitioning into the sensorless vector control can lead to an overshoot in the supply current to the motor and/or a jerk in the torque of the motor. [0034] In a different embodiment, the motor controller comprises a control circuit configured to implement and switch between the different control modes, and can detect the failure at the signal input, and is configured to determine or estimate the rotor position. The control circuit can be a digital processor, a microcontroller, an FPGA, or a similar means and can include analog components to convert the signals received at the signal input into a required digital format. The control circuit can also be configured to generate control signals for semiconductor switches comprised in a power stage of the motor controller. [0035] In a further embodiment, the motor controller comprises current sensors and voltage sensors, wherein each of the said sensors is configured to measure electrical quantities in the set of driving signals and of the electrical energy receivable at the output end. The electrical quantities correspond to voltages or currents comprised in the set of driving signals. [0036] In a further embodiment, the motor controller comprises a switching device configured to conduct or insulate the set of driving signals and of the electrical energy receivable at the output end. The motor H55-24-PCT controller can furthermore be configured to switchable connect the electrical source at the input end. Preferably the motor controller can be configured with a direct current link, comprising multiple capacitors. [0037] The switching device can be a contactor that is suitable to conduct or interrupt the power that is supplied to the motor and thus can be configured as a high power, high current contactor. Suppose the motor controller detects a failure at the signal input, e.g. caused by the defective speed sensor, the motor controller can disconnect the motor from the output end. [0038] With the use of voltage sensors connected to the output end, the phase, frequency and/or amplitude of the voltage that can be supplied by the motor can be determined to determine the virtual rotor position. The properties of the voltage measurable in this situation relate to the open- circuit characteristics of the motor. [0039] The electrical source can also be disconnected from the motor controller. Also, in this case, the motor can supply energy to the motor controller, and the motor controller or the control circuit can determine the rotor position. The energy supplied by the motor can be apparent in the direct current link, whereas the direct current link can be permanently charged with said energy. [0040] Some or all previously embodiments can be combined when it is useful and feasible from a technical standpoint. [0041] According to another aspect, a control system for an electric or hybrid aircraft is disclosed, which comprises: - an electrical source; - a first motor having a rotor; - a speed or position sensor coupled to the rotor for measuring a speed and/or a position of the said rotor; - a motor controller as previously described (including any embodiments or H55-24-PCT any combination thereof) connected to the electrical source at the input end, connected to the speed or position sensor at the signal input, and connected to the first motor at the output end. [0042] In an embodiment, the first motor is configured as a synchronous machine comprising a propeller coupled to the rotor. The synchronous machine can be a three-phase synchronous motor in the form of a three- phase permanent-magnet motor or a three-phase shunt motor. The propeller can be configured with a plurality of propeller blades, wherein the pitch or attack angle of said the blades can be varied to adjust the thrust. The propeller is conventionally mechanically connected to the motor shaft. [0043] In another embodiment, the control system comprises a second motor mechanically coupled to the rotor of the first motor. The second motor can be configured with a smaller power rating compared to the first motor. The second motor can be energized only when a failure at the signal input is detected, which may be caused by the loss or the failure of the sensor. [0044] The second motor can be used to keep the first motor rotating, such that the first motor supplies energy to the output end of the motor controller, for determining the rotor position of the first motor. For instance, if the second control scheme is entered, the second motor can be de-energized. The second motor can be useful if the aircraft is grounded and there is no airstream surrounding the propeller for turning the rotor of the first motor to supply electrical energy to the motor controller. [0045] According to another aspect, a method for operating the above described control system is disclosed, which comprises the steps of: - optionally controlling the rotor of the first motor from standstill to an operational speed by controlling the set of driving signals of the motor controller using the sensor-based control mode or the sensorless control mode; H55-24-PCT - detecting the failure at the signal input; - switching the control mode from the sensor-based or sensorless control mode to the initialization control mode, such that a remaining rotation of the rotor of the first motor supplies the motor controller at the output end with electrical energy; - determining or estimating the rotor position of rotor of the first motor from the electrical energy received at the output end; - switching the control mode from the initialization control mode to the sensorless control mode; - controlling the set of driving signal at the output end of the motor controller using the determined or estimated rotor position and the sensorless vector control. [0046] A failure of the speed or position sensor can thus be detected, and the control method of the motor controller can be adapted accordingly to cope with the potential loss of the sensor. The operation of the motor controller and finally of the motor connected to the motor controller can be continued, which leads to a significant improvement in the operability of the aircraft, in particular after a suspected sensor failure. The risk for catastrophic scenarios, such as outlined in the introduction, can be reduced by implementing the method on a motor controller. [0047] In an embodiment, the method comprises the step of maintaining the operational speed of the rotor by controlling the set of driving signal at the output end of the motor controller using the sensorless vector control. By using the sensorless vector control, the impact of the loss of the sensor signal can be reduced. [0048] The steps can be performed in the sequence as listed, or in any different sequence. [0049] In a further embodiment, the method comprises the step of energizing the second motor during an activation of the initialization control mode, for rotating the rotor of the first motor. This step can H55-24-PCT comprise exclusively energizing the second motor and thereby not energizing the first motor. By rotating the rotor of the first motor with the use of the second motor, the first motor can supply electrical energy to the motor controller, wherein the motor controller can in return determine the rotor position of the rotor of the first motor. [0050] In another embodiment, the method comprises the step of changing a blade pitch of the propeller coupled to the rotor during an activation of the initialization control mode for maintaining the operational speed of the rotor. Adapting the blade pitch can also be useful to increase the speed of the rotor, in particular for supplying sufficient energy to the motor controller to determine the rotor position. The blade pitch can preferably be adjusted during the flight of the aircraft, in particular when the blades are surrounded by the airstream. However, the blade pitch can also be adjusted during the operation on the ground in particular when the first motor is started using the V/f- or I/f-control, such that the required torque for the motor can be reduced. [0051] In a further embodiment, the method comprises the step of decoupling the first motor from the motor controller by insulating the set of driving signals with the use of the switching device during an activation of the initialization control mode. By insulating or decoupling the first motor from the motor controller the no-load voltage of the first motor can be measured at the output end, with the use of the voltage sensors, and a related rotor position can be determined by the motor controller. [0052] An electric or hybrid airplane propelled by the above described control system is also disclosed. [0053] With the use of the method, the motor controller can benefit from the airstream surrounding the propeller of the motor, so that the rotor position can be efficiently determined for a smooth transition between the sensor-based operation mode and the sensorless operation H55-24-PCT mode, without negative affecting the comfort or the safety of passengers in the airplane. Short description of the drawings [0054] Exemplar embodiments are disclosed in the description and illustrated by the drawings in which: Fig.1A illustrates an aircraft, such as an electric or hybrid aircraft; Fig.1B illustrates a simplified block diagram of an aircraft; Fig.2 illustrates management systems for operating an aircraft; Fig.3 illustrates a battery monitoring system for an aircraft; Figs.4 and 5 illustrate implementations of battery monitoring circuits; Figs.6 and 7 illustrate implementations of master circuits for monitoring battery monitoring circuits; Figs.8, 9, 10, 11, 12, and 13 illustrate schematic views of implementations of a power management system; Figs.14A and 14B illustrate a battery module usable in an aircraft; Figs.15A and 15B illustrate a power source formed of multiple battery modules; Fig.16 illustrates multiple power sources arranged and connected for powering an aircraft; H55-24-PCT Figs.17A and 17B illustrate multiple power sources positioned in a nose of an aircraft for powering the aircraft; Figs.18A and 18B illustrate multiple power sources positioned in a wing of an aircraft for powering the aircraft; Fig.19 illustrates a motor with multiple field coils; Figs.20 and 21 illustrate motors connected to a motor controller in different arrangements; Fig.22 illustrates a simplified control diagram for a motor controller connected to a motor; Examples of embodiments System Overview [0055] Fig.1A illustrates an aircraft 100, such as an electric or hybrid aircraft, and Fig.1B illustrates a simplified block diagram of the aircraft 100. The aircraft 100 includes a motor 110, a management system 120, and a power source 130. The motor 110 can be used to propel the aircraft 100 and cause the aircraft 100 to fly and navigate. The management system 120 can control and monitor the components (equipment) of the aircraft 100, such as the motor 110 and the power source 130. The power source 130 can power the motor 110 to drive the aircraft 100 and power the management system 120 to enable operations of the management system 120. The management system 120 can include one or more motor controllers as well as other electronic circuitry for controlling and monitoring various components of the aircraft 100. [0056] Fig.2 illustrates components 200 of an aircraft, such as the aircraft 100 of Figs.1A and 1B. The components 200 can include a power management system 210, a motor management system 220, and a recorder H55-24-PCT 230, as well as a first battery pack 212A, a second battery pack 212B, a warning panel 214, a fuse and relay 216, a converter 217, a cockpit battery pack 218, a motor controller 222, one or more motors 224, and a throttle 226. [0057] The power management system 210, the motor management system 220, and the recorder 230 can monitor communications on a communication bus, such as a controller area network (CAN) bus, and communicate via the communication bus. The first battery pack 212A and the second battery pack 212B can, for instance, communicate on the communication bus enabling the power management system 210 to monitor and control the first battery pack 212A and the second battery pack 212B. As another example, the motor controller 222 can communicate on the communication bus enabling the motor management system 220 to monitor and control the motor controller 222. [0058] The recorder 230 can store some or all data communicated (such as component status, temperature, or over/undervoltage information from the components or other sensors) on the communication bus to a memory device for later reference, such as for reference by the power management system 210 or the motor management system 220 or for use in troubleshooting or debugging by a maintenance worker. The power management system 210 and the motor management system 220 can each output or include a user interface that presents status information and permits system configurations. The power management system 210 can control a charging process (for instance, a charge timing, current level, or voltage level) for the aircraft when the aircraft is coupled to an external power source to charge a power source of the aircraft, such as the first battery pack 212A or the second battery pack 212B. [0059] The warning panel 214 can be a panel that alerts a pilot or another individual or computer to an issue, such as a problem associated with a power source like the first battery pack 212A. The fuse and relay 216 can be associated with the first battery pack 212A and the second H55-24-PCT battery pack 212B and usable to transfer power through a converter 217 (for example, a DC-DC converter) to a cockpit battery pack 218. The fuse and relay 216 can protect one or more battery poles of the first battery pack 212A and the second battery pack 212B from a short or overcurrent. The cockpit battery pack 218 may supply power for the communication bus. [0060] The motor management system 220 can provide control commands to the motor controller 222, which can in turn be used to operate the one or more motors 224. The motor controller can include an inverter for generating AC currents that are needed for operating the one or more motors. The motor controller 222 may further operate according to instructions from the throttle 226 that may be controlled by a pilot of the aircraft. The one or more motors can include an electric brushless motor. [0061] The power management system 210 and the motor management system 220 can execute the same or similar software instructions and may perform the same or similar functions as one another. The power management system 210, however, may be primarily responsible for power management functions while the motor management system 220 may be secondarily responsible for the power management functions. Similarly, the motor management system 220 may be primarily responsible for motor management functions while the power management system 210 may be secondarily responsible for the motor management functions. The power management system 210 and the motor management system 220 can be assigned respective functions, for example, according to system configurations, such as one or more memory flags in memory that indicate a desired functionality. The power management system 210 and the motor management system 220 may include the same or similar computer hardware. [0062] The power management system 210 can automatically perform the motor management functions when the motor management system 220 is not operational (such as in the event of a rebooting or failure of the motor management system 220), and the motor management system 220 H55-24-PCT can automatically perform the power management functions when the power management system 210 is not operational (such as in the event of rebooting or failure of the power management system 210). Moreover, the power management system 210 and the motor management system 220 can take over the functions from one another without communicating operation data, such as data about one or more of the components being controlled or monitored by the power management system 210 and the motor management system 220. This can be because both the power management system 210 and the motor management system 220 may be consistently monitoring communications on the communication bus to generate control information, but the control information may be used if the power management system 210 and the motor management system 220 has primary responsibility but not if the power management system 210 and the motor management system 220 does not have primary responsibility. Additionally or alternatively, the power management system 210 and the motor management system 220 may access data stored by the recorder 230 to obtain information usable to take over primary responsibility. System Architecture [0063] Electric and hybrid aircraft (rather than aircraft powered during operation by combustion) have been designed and manufactured for decades. However, electric and hybrid aircraft have still not yet become widely used for most transport applications like carrying passengers or goods. [0064] This failure to adopt may be in large part because designing an aircraft that is sufficiently safe to be certified by certification authorities may be very difficult. The certification of prototypes may moreover not be sufficient to certify for commercial applications. Instead, a certification of each individual aircraft and its components may be required. H55-24-PCT [0065] This disclosure provides at least some approaches for constructing electric powered aircraft from components and systems that have been designed to pass certification requirements so that the aircraft itself may pass certification requirements and proceed to active commercial use. [0066] Certification requirements can be related to a safety risk analysis. A condition that may occur with an aircraft or its components can be assigned to one of multiple safety risk assessments, which may in turn be associated with a particular certification standard. The condition can, for example, be catastrophic, hazardous, major, minor, or no safety effect. A catastrophic condition may be one that likely results in multiple fatalities or loss of the aircraft. A hazardous condition may reduce the capability of the aircraft or the operator ability to cope with adverse conditions to the extent that there would be a large reduction in safety margin or functional capability crew physical distress/excessive workload such that operators cannot be relied upon to perform required tasks accurately or completely or serious or fatal injury to small number of occupants of aircraft (except operators) or fatal injury to ground personnel or general public. A major condition can reduce the capability of the aircraft or the operators to cope with adverse operating condition to the extent that there would be a significant reduction in safety margin or functional capability, significant increase in operator workload, conditions impairing operator efficiency or creating significant discomfort physical distress to occupants of aircraft (except operator), which can include injuries, major occupational illness, major environmental damage, or major property damage. A minor condition may not significantly reduce system safety such that actions required by operators are well within their capabilities and may include a slight reduction in safety margin or functional capabilities, slight increase in workload such as routine flight plan changes, some physical discomfort to occupants or aircraft (except operators), minor occupational illness, minor environmental damage, or minor property damage. A no safety effect condition may be one that has not effect on safety. [0067] An aircraft can be designed so that different monitoring and warning subsystems, such as battery monitoring circuits, of the aircraft are H55-24-PCT constructed to have a robustness corresponding to their responsibilities and any related certification standards, as well as potentially any subsystem redundancies. [0068] Where a potential failure of the responsibilities of a monitoring and warning subsystem would likely be catastrophic, the subsystem can be designed to be simple and robust and thus may be able to satisfy difficult certification standards. The subsystem, for instance a battery, motor or motor controller monitoring circuit, can be composed of non- programmable, non-stateful components (for example, analog or non- programmable combinational logic electronic components) rather than programmable components (for example, a processor, a field programmable gate array (FPGA), or a complex programmable logic device (CPLD)) or stateful components (for example, sequential logic electronic components) and activate indicators such as lights rather than more sophisticated displays. [0069] On the other hand, where either (i) a monitoring and warning subsystem (such as a battery monitoring circuit, a motor monitoring circuit or a motor controller monitoring circuit) of an aircraft monitors a parameter redundantly with another subsystem of the aircraft that is composed of non-programmable, non-stateful components or (ii) a potential failure of the responsibilities of such a monitoring and warning subsystem would likely be less than catastrophic, or less than hazardous, the subsystem can be at least partly digital and designed to be complicated, feature-rich, and easier to update and yet able to satisfy associated certification standards. Such a subsystem can, for instance, include a processor or other programmable components that outputs information to a sophisticated display for presentation. [0070] In some implementations, some or all catastrophic conditions monitored for by an aircraft can be monitored for with at least one monitoring and warning subsystem that does not include a programmable component or a stateful component because certifications for H55-24-PCT programmable components or stateful components may demand statistical analysis of the responsible subsystems, which can be very expensive and complicated to certify. Such implementations can moreover be counterintuitive at least because an electric or hybrid aircraft may include one or more relatively advanced programmable or stateful components to enable operation of the electric or hybrid aircraft, so the inclusion of one or more subsystems in the aircraft that does not include any programmable components or any stateful components may be unexpected because the one or more relatively advanced programmable or stateful components may be readily and easily able to implement the functionality of the one or more subsystems that does not include any programmable components or any stateful components. [0071] An aircraft monitoring system can include a first monitoring and warning subsystem and a second monitoring and warning subsystem. The second subsystem, such as a second battery monitoring circuit, can be supported by an aircraft housing and include non-programmable, non- stateful components, such as analog or non-programmable combinational logic electronic components. The non-programmable, non-stateful components can monitor a component (such as battery cells in a battery pack) supported by the aircraft housing and output a second alert to notify of a catastrophic condition associated with the component. The non- programmable, non-stateful components can, for instance, activate an indicator or an audible alarm for a passenger aboard the housing to output the first alert. The indicator or audible alarm may remain inactive unless the indicator is outputting the first alert. Additionally or alternatively, the non-programmable, non-stateful components can output the second alert to a computer aboard or remote from the aircraft (for example, to automatically trigger actions to attempt to respond to or address the catastrophic condition, such as to stop charging or activate a fire extinguisher, a parachute, or an emergency landing procedure or other emergency response feature) or an operator of the aircraft via a telemetry system. The non-programmable, non-stateful components may, moreover, not be able to control the component or at least control certain H55-24-PCT functionality of the component, such as to control a mode or trigger an operation of the component. [0072] The first subsystem, such as a first battery monitoring circuit, can be supported by the aircraft housing and include a processor (or another programmable or stateful component), as well as a communication bus. The processor can monitor the component from communications on the communication bus and output a first alert to notify of a catastrophic condition or a less than catastrophic condition associated with the component. The processor can, for instance, activate an indicator or audible alarm for a passenger aboard the housing to output the first alert. Additionally or alternatively, the processor can output the first alert to a computer aboard or remote from the aircraft (for example, to automatically trigger actions to attempt to address the catastrophic condition, such as to activate a fire extinguisher, a parachute, or an emergency landing procedure) or an operator of the aircraft via a telemetry system. The processor may control the component. [0073] The non-programmable, non-stateful components of the second subsystem additionally may not be able to communicate via the communication bus. It may not include any programmable communication circuit for allowing communication via such a bus. [0074] An example of such a design and its benefits are next described in the context of battery management systems. Notably, the design can be additionally or alternatively applied to other systems of a vehicle that perform functions other than battery management, such as motor and motor control.

[0075] Battery packs including multiple battery cells, such as lithium-ion cells, can be used in electric cars, electric aircraft, and other electric self- H55-24-PCT powered vehicles. The battery cells may be connected in series or in parallel to deliver an appropriate voltage and current. [0076] Battery cells in battery packs can be managed and controlled by battery management systems (BMS). A BMS can be a circuit that manages a rechargeable battery cell by controlling its charging and discharge cycles, preventing it from operating outside its safe operating area, balancing the charge between cells, or the like. BMS can also monitor battery parameters, such as the temperature, voltage, current, internal resistance, or pressure of the battery cell, and report anomalies. BMS can be provided by various manufacturers as discreet electronic components. [0077] Damage to battery cells can be very serious incidents that may cause fire, explosions, or interruption of the powered circuit. Therefore, any damage to a battery in a vehicle, such as an electric airplane, may desirably be reported immediately and reliably to the pilot or driver of the vehicle. A reliable monitoring of battery cells by BMS can be critical for the safety of electric airplanes. [0078] However, BMS can have failings in rare occurrences that cause problems with battery cells which may not be reported correctly. For example, an overvoltage or overtemperature condition can, in some situations, affect not only a battery cell, but also its BMS, so that the failure of the battery cell is either not detected or not reported correctly. Even if the BMS functions correctly, a connecting bus between the BMS and the Cockpit might be defective and prevent warning signals from being transmitted. [0079] In order to prevent this risk, battery cells can be monitored with a second, redundant BMS. If both BMS are of the same type, a defect or conception flaw that affects one BMS may also affect the redundant BMS as well, so that the gain in reliability can be limited. The present disclosure provides at least approaches to increase the reliability of the detection of malfunctions of battery cells in an electric vehicle, such as an electric H55-24-PCT aircraft. Redundant monitoring of parameters of each battery cell can be performed with two different circuits. Because a second, redundant monitoring circuit may include non-programmable, non-stateful components rather than processors, sequential logic electronic components, or programmable combinational logic electronic components, its certification can be easier, and its reliability may be increased. For example, because the second, redundant circuit may be processorless, may not include any sequential or programmable combinational logic electronic components, and may not rely on any software (for example, executable program code that is executed by a processor), its certification is made easier than if the second, redundant circuit relied on processors, sequential or programmable combinational logic electronic components, or software. [0080] The second, redundant monitoring circuit can provide for a redundant monitoring of battery parameters and for a redundant transmission of those parameters, or warning signals depending on those parameters. The second battery monitoring system may transmit analog or binary signals but not multivalued digital signals. The second battery monitoring circuit may not manage the charge and discharge of battery cells, but instead provide for monitoring of battery parameters, and transmission of parameters or warning signals. Therefore, the second, redundant battery monitoring circuit can be made simple, easy to certify, and reliable. [0081] Fig.3 illustrates a battery monitoring system. This system can be used in an electric vehicle, such as an electric aircraft, a large size drone or unmanned aerial vehicle, an electric car, or the like, to monitor the state of battery cells 1 in one of multiple battery packs and report this state or generate warning signals in case of dysfunctions. [0082] The battery cells 1 can be connected in series or in parallel to deliver a desired voltage and current. Fig.3 shows serially connected battery cells. The total number of battery cells 1 may exceed 100 cells in an H55-24-PCT electric aircraft. Each of the battery cells 1 can be made up of multiple elementary battery cells in parallel. [0083] A first battery monitoring circuit can control and monitor the state of each battery cell 1. The first battery management circuit can include multiple BMSs 2, each of the BMSs 2 managing and controlling one of the battery cells 1. The BMSs 2 can each be made up of an integrated circuit (for instance, a dedicated integrated circuit) mounted on one printed circuit board (PCB) of the PCBs 20. One of the PCBs 20 can be used for each of the battery cells 1 or for a group of battery cells. Fig.4 illustrates example components of one of the BMSs 2. [0084] The control of a battery cell can include control of its charging and discharge cycles, preventing a battery cell from operating outside its safe operating area, or balancing the charge between different cells. [0085] The monitoring of one of the battery cells 1 by one of the BMSs 2 can include measuring parameters of the one of the battery cells 1, to detect and report its condition and possible dysfunctions. The measurement of the parameters can be performed with battery cell parameter sensors, which can be integrated in the one of the BMSs 2 or connected to the one of the BMSs 2. Examples of such parameter sensors can include a temperature sensor 21, a voltage sensor 22, or a current sensor. An analog- to-digital converter 23 can convert the analog values measured by one or more of the parameter sensors into multivalued digital values, for example, 8 or 16 bits digital parameter values. A microcontroller 24, which can be part of each of the BMSs 2, can compare the values with thresholds to detect when a battery cell temperature, battery cell voltage, or battery cell current is outside a range. [0086] The BMSs 2 as slaves can be controlled by one of multiple first master circuits 5. In the example of Fig.3, each of the first master circuits 5 can control four of the BMSs 2. Each of the first master circuits 5 can control eight of the BMSs 2, or more than eight of the BMSs 2. The first master H55-24-PCT circuits 5 can control more BMS and more battery cells in yet other implementations. The first master circuits 5 can be connected and communicate over a digital communication bus 55. [0087] The first master circuits 5 can also be connected to a computer 9 that collects the various digital signals and data sent by the first master circuits 5, and may display information related to the battery state and warning signals on a display 13, such as a matrix display. The display 13 may be mounted in the vehicle’s cockpit to be visible by the vehicle’s driver or pilot. Additionally or alternatively, the computer 9 can output the information to a computer remote from the aircraft or to control operations of one or more components of the aircraft as described herein. [0088] The BMSs 2 can be connected to the first master circuits 5 over a digital communication bus, such as a CAN bus. A bus driver 25 can interface the microcontroller 24 with the digital communication bus and provide a first galvanic isolation 59 between the PCBs 20 and the first master circuits 5. In one example, the bus drivers of adjacent BMSs 2 can be daisy chained. For example, as shown in Fig.4, the bus driver 25 is connected to the bus driver 27 of the previous BMS and to the bus driver 28 of the next BMS. [0089] Each of the BMSs 2 and their associated microcontrollers can be rebooted by switching its power voltage Vcc. The interruption of Vcc can be controlled by the first master circuits 5 over the digital communication bus and a power source 26. [0090] Fig.6 illustrates example components of one of the first master circuits 5. The one of the first master circuits 5 can include a first driver 51 for connecting the one of the first master circuits 5 with one of the BMSs 2 over the digital communication bus, a microcontroller 50, and a second driver 52 for connecting the first master circuits 5 between themselves and with the computer 9 over a second digital communication bus 55, such as a second CAN bus. A second galvanic isolation 58 can be provided between the first and second master circuits 5, 7 and the computer 9.27ulfild H55-24-PCT galvanic isolation 58 can, for example, be 1500 VDC, 2500 Vrms, 3750 Vrms, or another magnitude of isolation. The microcontroller 50, the first driver 51, and the second driver 52 can be powered by a powering circuit 53 and may be mounted on a PCB 54, one such PCB can be provided for each of the first master circuits 5. [0091] Fig.3 further illustrates a second battery monitoring circuit, which can be redundant of the first battery monitoring circuit. This second battery monitoring circuit may not manage the battery cells 1; for example, the second battery monitoring circuit may not control charge or discharge cycles of the battery cells 1. The function of the second battery monitoring circuit can instead be to provide a separate, redundant monitoring of each of the battery cells 1 in the battery packs, and to transmit those parameters or warning signals related to those parameters, such as to the pilot or driver or a computer aboard or remote from the aircraft as described herein. The second battery monitoring circuit can monitor the state of each of the battery cells 1 independently from the first battery monitoring circuit. The second battery monitoring circuit can include one of multiple cell monitoring circuits 3 for each of the battery cells. The parameters or warning signals may moreover, for example, be used by the second battery monitoring circuit to stop charging (for instance, by opening a relay to disconnect supply of power) of one or more battery cells when the one or more battery cells may be full of energy and a computer of the aircraft continues to charge the one or more battery cells. [0092] Fig.5 illustrates example components of one of the cell monitoring circuits 3. Each of the cell monitoring circuits 3 can include multiple cell parameter sensors 30, 31, 32, 33 for measuring various parameters of one of the battery cells 1. The sensor 30 can measure a first temperature at a first location in one battery cell and detect an overtemperature condition; the sensor 31 can measure a second temperature at a second location in the same battery cell and detect an overtemperature condition; the sensor 32 can detect an undervoltage condition in the same battery cell; and the sensor 33 can detect an overvoltage condition on the same battery cell. The undervoltage condition H55-24-PCT can be detected, for example, when the voltage at the output of one battery cell is under 3.1 Volts or another threshold. The overvoltage condition might be detected, for example, when the voltage at the output of one battery cell is above 4.2 Volts or yet another threshold. The thresholds used can depend, for instance, on the type of battery cell 1 or a number of elementary cells in the cell. Therefore, each or some of the sensors 30-33 can include a sensor as such and an analog comparator for comparing the value delivered by the sensor with one or two thresholds, and outputting a binary value depending on the result of the comparison. Other battery cell parameter sensors, such as an overcurrent detecting sensor, can be used in other implementations. [0093] Various parameters related to one of the battery cells 1 can be combined using a combinational logic circuit 35, such as an AND gate. The combinational logic circuit 35 may not include programmable logic. In the example of Fig.5, binary signals output by the sensors 30, 31, and 32 are combined by a AND gate into a single warning signal, which can have a positive value (warning signal) if and only if the temperature measured by the two temperature sensors exceeds a temperature threshold and if the voltage of the cell is under a voltage threshold. The detection of an overvoltage condition by the sensor 33, in the example of Fig.5, may not combined with any other measure and can be directly used as a warning signal. [0094] The warning signals delivered by the combinational logic 35 or directly by the parameter sensors 30-33 can be transmitted to a second master circuit 7 over lines 76, which can be dedicated and different from the digital communication bus used by the first battery monitoring circuit. Optocouplers 36, 37, 38 provide a third galvanic isolation 60 between the components 30-38 and the second master circuit 7. The third galvanic isolation 60 can provide the same isolation as the first galvanic isolation 59, such as 30V isolation, or the third galvanic isolation 60 may provide a different isolation form the first galvanic isolation 59. H55-24-PCT [0095] The sensors 30-33 and the combinational logic element 35 can be powered by a powering circuit 34 that delivers a power voltage Vcc2. This powering circuit 34 can be reset from the second master circuit 7 using an ON/OFF signal transmitted over the optocoupler 38. [0096] The sensors 30-33 and the combinational logic element 35 can be mounted on a PCB. One such PCB can be provided for each of the battery cells 1. The sensors 30-33 and the combinational logic element 35 can be mounted on the same PCB 20 as one of the BMSs 2 of the first battery monitoring circuit. [0097] Fig.7 illustrates example components of one of the second master circuits 7. In the example of Fig.5, the one of the second master circuits 7 can include a combinational logic element 72, which may not include programmable logic, for combining warning signals, such as overtemperature/under- voltage warning signals uv1, uv2, ... or overvoltage signals ov1, ov2, ... from different battery cells into combined warning signals, such as a general uv (undervoltage condition in case of overtemperature) warning signal and a separate overvoltage warning signal ov. Those warning signals uv, ov can be active when any of the battery cells 1 monitored by the one of the second master circuits 7 has a failure. They can be transmitted over optocouplers 70, 71 and lines 76 to the next and previous second master circuits 74, 75, and to a warning display panel 11 in the cockpit of the vehicle for displaying warning signals to the driver or pilot. The warning display panel 11 can include lights, such as light emitting diodes (LEDs), for displaying warning signals. [0098] With the disclosed design of the cell monitoring circuits 3 and the second master circuits 7, no dormant alarms may remain undetected. For example, if a cable may be broken or a power supply is inactive, the warning panel 11 can correctly show an alarm despite the broken cable or the inactive power supply. This can be accomplished, for instance, by using an inverted logic so that if the warning panel 11 does not receive a voltage or a current on an alarm line, an indicator may activate, but if the warning H55-24-PCT panel 11 does receive the voltage or the current on the alarm line, the indicator can deactivate. [0099] The one of the second master circuits 7 can be mounted on a PCB. One such PCB can be provided for each of the second master circuits 7. One of the second master circuits 7 can be mounted on the same PCB 54 as one of the first master circuits 5 of the first battery monitoring circuit. [00100] As can be seen, the second battery monitoring circuit can include exclusively non-programmable, non-stateful components (such as, analog components or non-programmable combinational logic components). The second battery monitoring circuit can be processorless, and may not include any sequential or programmable combinational logic. The second battery monitoring circuit may not run any computer code or be programmable. This simplicity can provide for a very reliable second monitoring circuit, and for a simple certification of the second battery monitoring circuit and an entire system that include the second battery monitoring circuit. [00101] The second battery monitoring circuit can be built so that any faulty line, components, or power source triggers an alarm. In one example, an “0” on a line, which may be caused by the detection of a problem in a cell or by a defective sensor, line, or electronic component, can be signalled as an alarm on the warning panel; the alarm may only be removed when all the monitored cells and all the monitoring components are functioning properly. For example, if the voltage comparator or temperature sensor is broken, an alarm can be triggered. [00102] The computer 9, the display 13, and the warning display panel 11 in the cockpit can be powered by a power source 15 in the cockpit, which may be a cockpit battery and can be independent of other power sources used to power one or more other components. H55-24-PCT Monitoring and warning about failure conditions in motors or motor controllers for electric and hybrid airplanes [00103] As indicated, the aspects, blocks and circuits that have been described so far in the context of battery monitoring systems could apply to monitoring and warning for different components of an electric or hybrid airplane. For example, a first monitoring and warning subsystem could be used for detecting an warning catastrophic, or hazardous, failure conditions of a motor or motor controller, while a second subsystem could be used for redundant monitoring of those catastrophic or hazardous failure conditions, and/or for monitoring and warning about less serious failure conditions, such as major, minor or no safety risk conditions of an electric motor or motor controller. The first monitoring and warning subsystem can be composed of non-programmable, non-stateful components and thus easier to certify, while the second monitoring and warning subsystem can comprise a processor or other programmable components, and output information to a sophisticated display 13, via a computer 9, for presentation. Motor and Battery System [00104] Battery packs including multiple battery cells, such as lithium-ion cells, can be used in electric cars, electric aircraft, and other electric self- powered vehicles. The battery cells can be connected in series or in parallel to deliver an appropriate voltage and current. [00105] In electrically driven aircraft, the battery packs can be chosen to fulfil the electrical requirements for various flight modes. During short time periods like take off, the electrical motor can utilize a relatively high power. During most of the time, such as in the standard flight mode, the electrical motor can utilize a relatively lower power, but may consume a high energy for achieving long distances of travel. It can be difficult for a single battery to achieve these two power utilizations. H55-24-PCT [00106] The use of two battery packs with different power or energy characteristics can optimize the use of the stored energy for different flight conditions. For example, a first battery pack can be used for standard flight situations, where high power output may not be demanded, but a high energy output may be demanded. A second battery pack can be used, alone or in addition to the first battery pack, for flight situations with high power output demands, such as take-off manoeuvring. [00107] An electrical powering system can charge the second battery pack from the first battery pack. This can allow recharging of the second battery pack during the flight, subsequent to the second battery pack being used in a high power output demanding flight situation. Therefore, the second battery pack can be small, which can save space and weight. In addition, this can allow different battery packs for different flight situations that optimize the use of the battery packs. [00108] The electrical powering system can also charge the second battery pack by at least one motor which works as generator (the motor may also accordingly be referred to as a transducer). This can allow recharging of the second battery pack during the flight or after the second battery pack has been used in a high power output demanding flight situation. Therefore, the second battery pack can be small, which can save space and weight. In addition, the different battery packs can allow the recovery of braking energy. Braking energy during landing or sinking recovered by a generator motor can create high currents which may not be recovered by battery packs used for traveling long distances. By using a second battery pack suitable for receiving high power output in a short time, more braking energy can be recovered via the second battery pack than the first battery pack, for example. [00109] The electrical powering system can also include a third battery pack, which includes a supercapacitor. Because supercapacitors can receive and output large instantaneous power or high energy in a short duration of time, the third battery pack can further improve the electrical powering H55-24-PCT system in some instances. A supercapacitor may, for example, have a capacitance of 0.1 F, 0.5 F, 1 F, 5 F, 10 F, 50 F, 100 F, or greater or within a range defined by one of the preceding capacitance values. [00110] Figs.8 to 13 illustrate multiple electrical power systems. [00111] Fig.8 shows an electrical powering system that includes a first battery pack 91, a second battery pack 92, a circuit 90, and at least one motor 94. [00112] The first battery pack 91 and the second battery pack 92 can each store electrical energy for driving the at least one motor 94. The first battery pack 91 and the second battery pack 92 can have different electrical characteristics. The first battery pack 91 can have a higher energy capacity per kilogram than the second battery pack 92, and the first battery pack 91 can have a higher power capacity (watt hours) than the second battery pack 92. Moreover, the first battery pack 91 can have a lower maximum, nominal, or peak power than the second battery pack 92; the first battery pack 91 can have a lower maximum, nominal, or peak current than the second battery pack 92; or, the first battery pack 91 can have a lower maximum, nominal, or peak voltage than the second battery pack 92. More than one or even all of the mentioned electrical characteristics of the first battery pack 91 and the second battery pack 92 can be different. However, only one of the mentioned electrical characteristics may be different or at least one other characteristic than the mentioned electrical characteristics may be different. The first battery pack 91 and the second battery pack 92 can have the same electrical characteristics. [00113] The type or the material composition of the battery cells of the first battery pack 91 and the second battery pack 92 can be different. The type or the material composition of the battery cells of the first battery pack 91 and the second battery pack 92 can be the same, but an amount of copper or an arrangement of conductors can be different. In one example, the first battery pack 91 or the second battery pack 92 can be a lithium-ion H55-24-PCT (Li-ion) battery or a lithium-ion polymer (Li-Po) battery. The second battery pack 92 may include a supercapacitor (sometimes referred to as a supercap, ultracapacitor, or Goldcap). [00114] The first battery pack 91 can include relatively high energy- density battery cells that can store a high amount of watt-hours per kilogram. The first battery pack 91 can include low power battery cells. The first battery pack 91 can provide a DC voltage/current/power or can be connected by a (two phase or DC) power line with the circuit 90. [00115] The second battery pack 92 can include relatively low energy- density battery cells. The second battery pack 92 can include relatively high power battery cells. The second battery pack 92 can provide a DC voltage/current/power or is connected by a (two phase or DC) power line with the circuit 90. [00116] The first battery pack 91 can form an integrated unit of mechanically coupled battery modules or the first battery pack 91 may be an electrically connected first set of battery modules. Similarly, the second battery pack 92 can form an integrated unit of mechanically coupled battery modules or the second battery pack 92 may be an electrically connected second set of battery modules. Some or all of the battery modules of each of first battery pack 91 or the second battery pack 92 can be stored in one or more areas of a housing of an aircraft, such as a within a wing or nose of the aircraft. [00117] The first battery pack 91 can have a total energy capacity that exceeds a total energy capacity of the second battery pack 92. For example, a ratio of the total energy capacity of the first battery pack 91 and the total energy capacity of the second battery pack 92 can be 2:1, 3:1, 4:1, 5:1, 10:1, 20:1, 40:1, or 100:1 or within a range defined by two of the foregoing ratios. H55-24-PCT [00118] The electrical powering system can include an external charging interface for charging the first battery pack 91 or the second battery pack 92 when the aircraft is on the ground and connected to a charging station outside of the aircraft. [00119] Each, some, or one of the at least one motor can be an electrical motor. The at least one motor 94 can be connected to the circuit 90. The at least one motor 94 can receive over the circuit 90 electrical energy/power from the first battery pack 91 or the second battery pack 92 to drive the at least one motor 94. For example, the at least one motor 94 can be a three phase motor, such as a brushless motor, which is connected via a three phase AC power line with the circuit 90. However, the at least one motor 94 can instead be a different type of motor, such as any type of DC motor or a one phase AC motor. The at least one motor 94 can move a vehicle, such as an airborne vehicle like an aircraft. The at least one motor 94 can drive a (thrust-generating) propeller or a (lift-generating) rotor. In addition, the at least one motor 94 can also function as a generator. The electrical powering system or the at least one motor 94 can include two or more electrical motors as described further herein. [00120] The different motors of the at least one motor 94 can have the same or different characteristics. The at least one motor 94 can be a motor with a first set of windings connected with a first controller 96 and with a second set of windings connected with a second controller 97, as shown for example in Fig.12. This can allow use of the at least one motor 94 as generator and motor at the same time or to power the at least one motor 94 from the first controller 96 and the second controller 97. The at least one motor 94 can include a first motor 98 and a second motor 99 as shown for example in Figs.11 and 13. The first and the second motor 98 and 99 can be mechanically connected such that the rotors of the first and second motor 98 and 99 are mechanically coupled, for instance for powering both the same propeller or rotor (as shown in Figs.11 and 13). The first and the second motor 98 and 99 can, for example, drive the same axis which rotates the propeller or rotor. However, the first and second motor 98 and 99 may not be mechanically coupled and may drive two distinct propellers or H55-24-PCT rotors. The at least one motor 94 can include more than two motors M1, M2, … Mi which are mutually connected, or multiple mutually connected motors. [00121] The circuit 90 can be connected with the first battery pack 91, the second battery pack 92, and the at least one motor 94. [00122] The circuit 90 can include a controller 93 connected with the first battery pack 91, the second battery pack 92, and the at least one motor 94. The controller 93 can, for example, be connected over a two phase or DC power line with the first battery pack 91 and the second battery pack 92 or connected over a three phase power line with the at least one motor 94. The controller 93 can transform, convert, or control the power received from the first battery pack 91 or the second battery pack 92 into motor driving signals for driving the at least one motor 94. The controller 93 can include a power converter for converting the DC current of the first battery pack 91 or the second battery pack 92 into a (three phase) (AC) current for the at least one motor 94 (power converter working as inverter). The power converter can treat different input DC voltages (if the first battery pack 91 and the second battery pack 92 have different DC voltages). If the at least one motor 94 acts as generator, the power converter can convert the current generated from each phase of the at least one motor 94 into a DC current for loading the first battery pack 91 or the battery pack 92 (power converter working as rectifier). The controller 93 can create the motor driving signals for the at least one motor 94 based on user input. [00123] The controller 93 can include more than one controller. The controller 93 can include, for instance, a first controller 96 for powering the at least one motor 94 from at least one of the first battery pack 91 and the second battery pack 92 and a second controller 97 for powering the at least one motor 94 from at least one of the first battery pack 91 or the second battery pack 92. The features described for the controller 93 can apply to the first controller 96 or the second controller 97. Examples of such a circuit are shown in the Figs.10 to 13. In Figs.10 to 12, the first controller 96 H55-24-PCT powers the at least one motor 94 from the first battery pack 91 and the second controller 97 powers the at least one motor 94 from the second battery pack 92. The first controller 96 and the second controller 97 can power the at least one motor 94 as shown in Fig.10 or the at least one motor 94 with different driving windings (or poles) as shown in Fig.12. [00124] As shown in Figs.11 and 13, the first controller 96 can drive a first motor 98 and the second controller 97 can drive a second motor 99. The first controller 96 and the second controller 97 can be flexible and drive the first motor 98 or the second motor 99 depending on a switching state of a switch 101 as shown in Fig.13. The first controller 96 and the second controller 97 can be different. For example, the input DC voltage of the first controller 96 and the second controller 97 from the first battery pack 91 and the second battery pack 92 can be different. However, the first controller 96 and the second controller 97 can instead be identical. [00125] The circuit 90 can select from at least two of the following connection modes. In a first connection mode, the first battery pack 91 can be electrically connected over the controller 93 with the at least one motor 94, while the second battery pack 92 may be electrically disconnected from the at least one motor 94. In the first connection mode, power can flow between the at least one motor 94 and the first battery pack 91, but may not flow between the at least one motor 94 and the second battery pack 92. In a second connection mode, the second battery pack 92 can be electrically connected over the controller 93 with the at least one motor 94, while the first battery pack 91 may be electrically disconnected from the at least one motor 94. In the second connection mode, power can flow between the at least one motor 94 and the second battery pack 92, but may not between the at least one motor 94 and the first battery pack 91. In a third connection mode, the first battery pack 91 and the second battery pack 92 can be electrically connected over the controller 93 with the at least one motor 94. In the third connection mode, power can flow between the at least one motor 94 and the first battery pack 91 and the second battery pack 92. Electrical switches can be used to perform this selection between different connection modes, and the electrical switches can be H55-24-PCT between the controller 93 and first battery pack 91 and the second battery pack 92, in the controller 93, or between the controller 93 and the at least one motor 94. If the at least one motor 94 has more than one motor, there can be further connection modes. The first battery pack 91 can be connected with the first motor 98 and not the second motor 99 (fourth connection mode) or with the second motor 99 and not the first motor 98 (fifth connection mode) or with the first motor 98 and the second motor 99 (sixth connection mode). The second battery pack 92 can be connected with the first motor 98 and not the second motor 99 (seventh connection mode) or with the second motor 99 and not the first motor 98 (eighth connection mode) or with the first motor 98 and the second motor 99 (ninth connection state). The first battery pack 91 and the second battery pack 92 can be connected with the first motor 98 and not the second motor 99 (tenth connection mode) or with the second motor 99 and not the first motor 98 (eleventh connection mode) or with the first motor 98 and the second motor 99 (twelfth connection state). The numbering of the connection modes can be arbitrarily chosen. If there may additionally be a third battery pack, there can be correspondingly more potential connection modes between the at least one motor and the three battery packs. [00126] The circuit 90 can select from at least two of the following drive modes. In a first drive mode, the at least one motor 94 can be driven by the first battery pack 91 (without using the power of the second battery pack 92). In this first drive mode (which may be referred to as a standard drive mode), the circuit 90 can be in the first connection mode. Alternatively, in the first drive mode, the circuit 90 can also be in the third connection mode, while no power flows from the second battery pack 92 to the at least one motor 94. This standard drive mode can be used when the power consumption of the least one motor 94 may be low, such as during steady flight conditions, gliding flight, or landing of an aircraft. In a second drive mode (which may be referred to as a high energy drive mode), the at least one motor 94 can be driven by the second battery pack 92 (without using the power of the first battery pack 91). In this second drive mode, the circuit 90 can be in the second connection mode. Alternatively, in the second drive mode, the circuit 90 can also be in the third motor connection H55-24-PCT mode, while no power flows from the first battery pack 91 to the at least one motor 94. This second drive mode can be used when the power consumption of the at least one motor 94 may be high, such as during manoeuvring, climb flight, or take off. In a third drive mode (which may be referred to as a very high energy drive mode), the at least one motor 94 can be simultaneously driven by the first battery pack 91 and the second battery pack 92. In this third drive mode, the circuit 90 can be in the third connection mode. This third drive mode can be used when the power consumption of the least one motor 94 may be high, such as during manoeuvring, climb flight, or take off. [00127] The circuit 90 can include a detector for detecting the power requirements of a present flight mode. The detection can be performed from user input or sensor measurements, such as by measuring the current in the motor input line. The circuit 90 can select the drive mode or the connection mode based at least on the detection result of this detector. [00128] The selection between connection modes can depend at least on the charging level of the different battery packs. For example, a high- power battery pack can be used instead, or in addition to, a high energy- density battery pack when the charge of the high energy-density battery pack is low. [00129] The electrical powering systems of Figs.8 to 13 can be configured such that the second battery pack 92 can be charged from the first battery pack 91, such as via the circuit 90. Moreover, the electrical powering systems can be configured such that the second battery pack 92 can be charged from the first battery pack 91 while the first battery pack 91 powers or drives the at least one motor 94. [00130] In Figs.9 to 11, the circuit 90 can electrically connect the first battery pack 91 and the second battery pack 92 for charging. The connection can be steady or realized by a switch which switches between a first battery connection mode in which the first battery pack 91 and the H55-24-PCT second battery pack 92 are electrically connected and a second battery connection mode in which the first battery pack 91 and the second battery pack 92 are electrically disconnected. As explained further herein, the first battery connection mode can be realized by connecting the first battery pack 91 and the second battery pack 92 over a charging circuit 95 or over the controller 93 or over one or more other controllers. [00131] In Fig.9, the circuit 90 the charging circuit 95 for charging the second battery pack 92 from the first battery pack 91. The charging circuit 95 can control energy flow from the first battery pack 91 to the second battery pack 92 and may transfer the energy without transferring the energy through the controller 93. The charging circuit 95 can include a switch (not shown) for connecting the first battery pack 91 with the second battery pack 92 for charging. Such a switch may have the advantage that the charging process can be controlled by a user or by a microprocessor. For example, if the full power of the first battery pack 91 is desired to power the at least one motor 94, the process of charging the second battery pack 92 may automatically be interrupted. However, the charging circuit 95 can instead work switchless so that the process of charging automatically starts when a certain electrical parameter, like the voltage or capacitance of the second battery pack 92, falls below a certain threshold. [00132] If the voltage of the first battery pack 91 and the second battery pack 92 may be different, the charging circuit 95 can include a DC/DC converter for converting the DC voltage of the first battery pack 91 into the DC voltage of the second battery pack 92. The second battery pack 92 can be charged from the first battery pack 91 at the same time that the at least one motor 94 is driven by the first battery pack 91 or at a time that the at least one motor 94 is not powered, such as by the first battery pack 91. [00133] In Fig.10, the second battery pack 92 can be charged over the first controller 96 and the second controller 97. The first battery pack 91 can provide energy and power for the first controller 96, which can convert this energy and power into the electrical driving signals for the at least one H55-24-PCT motor 94. For charging the second battery pack 92, the electrical driving signals from the first controller 96 can be converted by the second controller 97 into the charging signal (DC voltage) for the second battery pack 92. The electrical driving signals for the at least one motor 94 from the first controller 96 can be used for charging the second battery pack 92 and for driving the at least one motor 94 at the same time. This can allow the second battery pack 92 to charge from the first battery pack 91 at the same time that the at least one motor 94 may be driven by the electrical driving signals from the first controller 96. The second battery pack 92 can however instead be charged by the electrical drive signals without powering the motor at the same time. [00134] Instead of or in addition to electrically connecting the first battery pack 91 with the second battery pack 92 for transferring electrical energy from the first battery pack 91 to the second battery pack 92, the first battery pack 91 can be mechanically connected with the second battery pack 92 for transferring mechanical energy to charge the second battery pack 92 from the first battery pack 91. [00135] In Fig.11, mechanical charging can be realized by driving the first motor 98 from the first battery pack 91 (over the first controller 96) and generating energy from the second motor 99 which is mechanically connected to the first motor 98 and working as generator. The energy generated by the second motor 99 can be used to charge the second battery pack 92 (by converting the generated motor signals of the second motor 99 via the second controller 97 into the charging signal (DC voltage) of the second battery pack 92). This can allow the second battery pack 92 to charge from the first battery pack 91 at the same time that the at least one motor 94 is driven by the energy from the first battery pack 91. [00136] In Fig.12, mechanical charging can be realized by driving the at least one motor 94 from the first battery pack 91 (such as over the first controller 96) with the first set of windings of the at least one motor 94 and generating energy from the at least one motor 94 over the second set H55-24-PCT of windings of the at least one motor 94 which can function as a generator. The energy generated by the second set of windings can be used to charge the second battery pack 92 by converting the generated motor signals of the at least one motor 94 via the second controller 97 into the charging signal (DC voltage) of the second battery pack 92. This can allow the second battery pack 92 to charge from the first battery pack 91 at the same time that the at least one motor 94 is driven by the energy from the first battery pack 91. Moreover, this can enable the second battery pack 92 to charge from the first battery pack 91 without utilizing separate circuitry, such as a DC/DC converter, which would increase a weight of the aircraft. [00137] Fig.13 shows a switch 101 which can select from different battery packs or connection modes as described herein. This can allow the first battery pack 91 to connect with the second battery pack 92 (first battery connection mode) to charge the second battery pack 92 directly from the first battery pack 91. This can allow the first battery pack 91 to connect with (i) one of the first controller 96 or the second controller 97, (ii) one of the first motor 98 or second motor 99 and the second battery pack 92 with the other of the first controller 96 or the second controller 97, or (iii) the first motor 98 and the second motor 99 to charge the second battery pack 92 mechanically. This can allow for selection of the first motor 98 or the second motor 99 to be driven by the first battery pack 91 or the second battery pack 92. [00138] The design of Fig.13 can give the flexibility to choose among electrical charging or mechanical charging. [00139] The second battery pack 92 can be charged by the at least one motor 94 which can work as a generator. When the at least one motor 94 may work as a generator, the generation can be driven by braking energy, such as during descent or landing of the aircraft. The second battery pack 92 can as a result recover energy without affecting the functioning of the first battery pack 91 for long distances. When the at least one motor 94 may work as a generator, the generation can be driven from the first H55-24-PCT battery pack 91 to charge the second battery pack 92. The second battery pack 92 can be charged by the at least one motor 94 working as a generator while the same motor or another motor of the at least one motor 94 can be driven by the energy from the first battery pack 91, such as for instance described with respect to Figs.11, 12, and 13. [00140] The electrical powering system can include a third battery pack (not shown). The second battery pack 92 and the third battery pack can have different electrical characteristics. The second battery pack 92 can, for instance, have a higher energy capacity than the third battery pack. The second battery pack 92 can have a higher energy density than the third battery pack. The second battery pack 92 can have a lower maximum, nominal, or peak power than the third battery pack. The second battery pack 92 can have a lower maximum, nominal, or peak current than the third battery pack. The second battery pack 92 can have a lower maximum, nominal, or peak voltage than the third battery pack. The type or the material composition of the battery cells of the second battery pack 92 and the third battery pack can be different or the same. The third battery pack can include a supercapacitor. The third battery pack can increase a maximum power that may be delivered or recovered by the electrical powering system. The power recovered by the at least one motor 94 acting as a generator from a braking action can, for example, immediately be recovered in the third battery pack up to a high recover power level. The third battery pack can be charged from the first battery pack 91 or the second battery pack 92, such as even while the at least one motor 94 may be driven from the power of the first battery pack 91 or the second battery pack 92. Modular Battery System [00141] The power sources in an electric or hybrid aircraft can be modular and distributed to optimize a weight distribution or select a center of gravity for the electric or hybrid aircraft, as well as maximize a use of space in the aircraft. Moreover, the batteries in an electric or hybrid aircraft can H55-24-PCT desirably be designed to be positioned in place of a combustion engine so that the aircraft can retain a similar shape or structure to a traditional combustion powered aircraft and yet may be powered by batteries. In such designs, the weight of the batteries can be distributed to match that of a combustion engine to enable the electric or hybrid aircraft to fly similarly to the traditional combustion powered aircraft. [00142] Fig.14A illustrates a battery module 1400 usable in an aircraft, such as the aircraft 100 of Figs.1A and 1B. The battery module 1400 can include a lower battery module housing 1410, a middle battery module housing 1420, an upper battery module housing 1430, and a multiple battery cells 1440. The multiple battery cells 1440 can together provide output power for the battery module 1400. The lower battery module housing 1410, the middle battery module housing 1420, or the upper battery module housing 1430 can include slots, such as slots 1422, that are usable to mechanically couple the lower battery module housing 1410, the middle battery module housing 1420, or the upper battery module housing 1430 to one another or to another battery module. Supports, such as supports 1424 (for example, pins or locks), can be placed in the slots to lock the lower battery module housing 1410, the middle battery module housing 1420, or the upper battery module housing 1430 to one another or to another battery module. [00143] The battery module 1400 can be constructed so that the battery module 1400 is evenly cooled by air. The multiple battery cells 1440 can include 16 total battery cells where the battery cells are each substantially shaped as a cylinder. The lower battery module housing 1410, the middle battery module housing 1420, or the upper battery module housing 1430 can be formed of or include plastic and, when coupled together, have an outer shape substantially shaped as a rectangular prism. The lower battery module housing 1410, the middle battery module housing 1420, or the upper battery module housing 1430 can together be designed to prevent a fire in the multiple battery cells 1440 from spreading outside of the battery module 1400. H55-24-PCT [00144] The battery module 1400 can have a length of L1, a width of W, and a height of H1. The length of L1, the width of W, or the height of H1 can each be 50 mm, 65 mm, 80 mm, 100 mm, 120 mm, 150 mm, 200 mm, 250 mm or within a range defined by two of the foregoing values or another value greater or less than the foregoing values. [00145] Fig.14B illustrates an exploded view of the battery module 1400 of Fig.14A. In the exploded view, a plate 1450 and a circuit board assembly 1460 of the battery module 1400 is shown. The plate 1450 can be copper and may electrically connect the multiple battery cells 1440 in parallel with one another. The plate 1450 may also distribute heat evenly across the multiple battery cells 1440 so that the multiple battery cells 1440 age at the same rate. The circuit board assembly 1460 may transfer power from or to the multiple battery cells 1440, as well as include one or more sensors for monitoring a voltage or a temperature of one or more battery cells of the multiple battery cells 1440. The circuit board assembly 1460 may or may not provide galvanic isolation to the battery module 1400 with respect to any components that may be electrically connected to the battery module 1400. Each of the multiple battery cells 1440 can have a height of H2, such as 30 mm, 50 mm, 65 mm, 80 mm, 100 mm, 120 mm, 150 mm or within a range defined by two of the foregoing values or another value greater or less than the foregoing values. [00146] Fig.15A illustrate a power source 1500A formed of multiple battery modules 1400 of Figs.14A and 14B. The multiple battery modules 1400 of the power source 1500A can be mechanically coupled to one another. A first side of one battery module 1400 can be mechanically coupled to a first side of another battery module 1400, and a second side of the one battery module 1400 that is opposite the first side can be mechanically coupled to a first side of yet another battery module 1400. The multiple battery modules 1400 of the power source 1500A can be electrically connected in series with one another. As illustrated in Fig.15A, the power source 1500A can include seven of the battery modules 1400 connected to one another. The power source 1500A may, for example, have a maximum power output between 1 kW and 60 kW during H55-24-PCT operation, a maximum voltage output between 10 V and 120 V during operation, or a maximum current output between 100 A and 500 A during operation. [00147] The power source 1500A can include a power source housing 1510 mechanically coupled to at least one of the battery modules. The power source housing 1510 can include an end cover 1512 that covers a side of the power source housing 1510. The power source housing 1510 can have a length of L2, such as 3 mm, 5 mm, 10 mm, 15 mm, 20 mm, 25 mm, 30 mm, 40 mm, 50 mm or within a range defined by two of the foregoing values or another value greater or less than the foregoing values. The width and the height of the power source housing 1510 can match the length of L1 and the width of W of the battery module 1400. [00148] The power source 1500A can include power source connectors 1520. The power source connectors 1520 can be used to electrically connect the power source 1500A to another power source, such as another of the power source 1500A. [00149] Fig.15B illustrates a power source 1500B that is similar to the power source 1500A of Fig.15A but with the end cover 1512 and the upper battery module housings 1430 of the battery modules 1400 removed. Because the end cover 1512 has been removed, a circuit board assembly 1514 of the power source 1500B is now exposed. The circuit board assembly 1514 can be electrically coupled to the battery modules 1400. The circuit board assembly 1514 can additionally provide galvanic isolation (for instance, 2500 Vrms) for the power source 1500B with respect to any components that may be electrically connected to the power source 1500B. The inclusion of galvanic isolation in this manner may, for instance, enable grouping of the battery modules 1400 together so that isolation may be provided to the grouping of the battery modules 1400 rather than individual modules of the battery modules 1400 or a subset of the battery modules 1400. Such an approach may reduce the costs of construction H55-24-PCT because isolation can be expensive, and a single isolation may be used for multiple of the battery modules 1400. [00150] Fig.16 illustrates a group 1600 of multiple power sources 1500A of Fig.15A arranged and connected for powering an aircraft, such as the aircraft 100 of Figs.1A and 1B. The multiple power sources 1500A of the group 1600 can be mechanically coupled to or stacked on one another. The multiple power sources 1500A of the group 1600 can be electrically connected in series or parallel with one another, such as by a first connector 1610 or a second connector 1620 that electrically connects the power source connectors 1520 of two of the multiple power sources 1500A. As illustrated in Fig.16, the group 1600 can include 10 power sources (for instance, arranged in a 5 row by 2 column configuration). In other examples, a group may include a fewer or greater number of power sources, such as 2, 3, 5, 7, 8, 12, 15, 17, 20, 25, 30, 35, or 40 power sources. [00151] The grouping of the multiple power sources 1500A to form the group 1600 or another different group may allow for flexible configurations of the multiple power sources 1500A to satisfy various space or power requirements. Moreover, the grouping of the multiple power sources 1500A to form the group 1600 or another different group may permit relatively easy or inexpensive replacement of one or more of the multiple power sources 1500A in the event of a failure or other issue. [00152] Fig.17A illustrates a perspective view of a nose 1700 of an aircraft, such as the aircraft 100 of Figs.1A and 1B, that includes multiple power sources 1710, such as multiple of the power source 1500A, for powering a motor 1720 that operates a propeller 1730 of the aircraft. The multiple power sources 1710 can be used to additionally or alternatively power other components of the aircraft. The multiple power sources 1710 can be sized and arranged to optimize a weight distribution and use of space around the nose 1700. The motor 1720 and the propeller 1730 can be attached to and supported by a frame of the aircraft by supports, which can be steel tubes, and connected by multiple fasteners, which be bolts H55-24-PCT with rubber shock absorbers. A firewall 1740 can provide barrier between the multiple power sources 1710 and the frame of the aircraft in the event of a first at the multiple power sources 1710. An enclosure composed of glass fiber, metal, or mineral composite can be around the multiple power sources 1710 to protect from water, coolant, or fire. [00153] Fig 17B illustrates a side view of the nose 1700 of Fig.17A. [00154] Fig.18A illustrates a top view of a wing 1800 of an aircraft that includes multiple power sources 1810, such as multiple of the power source 1500A, for powering one or more components of the aircraft. The multiple power sources 1810 can be sized and arranged to optimize a weight distribution and use of space around the wing 1800. For example, the multiple power sources 1810 can be positioned within, between, or around horizontal support beams 1820 or vertical support beams 1830 of the wing 1800. A relay 1840 can further be positioned in the wing 1800 as illustrated and housed in a sealed enclosure. The relay 1840 may open if there is not a threshold voltage on a breaker panel or if a pilot opens breakers to shut down the multiple power sources 1810. [00155] Fig 18B illustrates a perspective view of the wing 1800 of Fig. 18A. Multi-Coil Motor Control [00156] An electric or hybrid aircraft can be powered by a multi-coil motor, such as an electric motor, in which different coils of the motor power different phases of a modulation cycle for the motor. [00157] As can be seen from Fig.19, a motor 1910 can include four different field coils (sometimes also referred to as coils) for generating a torque on a rotor of the motor 1910. The different field coils can include a first field coil 1902, a second field coil 1904, a third field coil 1906, and a H55-24-PCT fourth field coil 1908. Each of the different field coils can be independently powered by one or more controllers. The first field coil 1902, the second field coil 1904, the third field coil 1906, and the fourth field coil 1908 can be respectively powered by a first controller 1912, a second controller 1914, a third controller 1916, and a fourth controller 1918. One or more of the first controller 1912, the second controller 1914, the third controller 1916, and the fourth controller 1918 may be the same controller. [00158] The first controller 1912, the second controller 1914, the third controller 1916, and the fourth controller 1918 can vary a current provided to individual coils of the first field coil 1902, the second field coil 1904, the third field coil 1906, and the fourth field coil 1908 to compensate for a failure of one or more (such as, one, two, or three) of the field coils. The first controller 1912, the second controller 1914, the third controller 1916, and the fourth controller 1918 may, for example, no longer provide current to a coil that has failed and provide additional current to one or more coils that have not yet failed. The first controller 1912, the second controller 1914, the third controller 1916, and the fourth controller 1918 can attempt to maintain a power output of the motor (for example, above a threshold) despite the failure of the one or more of the field coils. [00159] The first controller 1912, the second controller 1914, the third controller 1916, or the fourth controller 1918 can determine the failure of one or more of the field coils from one or more sensors monitoring the motor or one or more individual field coils, such as proximate to the motor or one or more individual field coils. The one or more sensors can include a temperature sensor, a current sensor, or a magnetic field sensor, among other types of sensors. For example, where the one or more sensors includes at least one temperature sensor, the first controller 1912, the second controller 1914, the third controller 1916, or the fourth controller 1918 can determine the failure of one or more of the field coils from a change in the temperature sensed by the temperature sensor (for instance, a temperature drop over time or proximate different field coils may correspond to a failure of a particular field coil or a number of field coils in the motor 1910). The first controller 1912, the second controller 1914, the H55-24-PCT third controller 1916, or the fourth controller 1918 may moreover attempt to operate the motor so that the temperature sensed remains constant within a tolerance. As another example, where the one or more sensors includes at least one voltage sensor, the first controller 1912, the second controller 1914, the third controller 1916, or the fourth controller 1918 can determine the failure of one or more of the field coils from a change in the voltage sensed by the voltage sensor (for instance, a voltage spike may correspond to a failure of a particular field coil or a number of field coils in the motor 1910). In another example, where the one or more sensors includes at least one magnetic field sensor, the first controller 1912, the second controller 1914, the third controller 1916, or the fourth controller 1918 can determine the failure of one or more of the field coils from a change in the resonance sensed by the magnetic field sensor.

of the motor controller, the control system and the method for controlling the control

[00160] Figs.20 and 21 illustrate a motor 94 connected to the motor controller 93 in different arrangements. [00161] Fig.20 illustrates a first battery pack 91 and a second battery pack 92 connected in series for providing a DC voltage Vdc at the input side of the motor controller 93. The motor controller 93 supplies the motor 94 on the output side with electrical energy. [00162] The motor 94 in the following examples is configured as a three- phase permanent-magnet synchronous machine. Other motor types, such as induction machines, might be used instead in all examples. The motor 94 can propel the aircraft with the use of the propeller. [00163] The motor controller 93 provides the motor 94 with three-phase alternating currents and voltages with varying amplitudes and varying frequencies. The motor controller 93 is also configured to vary the fundamental frequency of the alternating currents and voltages provided H55-24-PCT to the motor 94. A propeller is attached to the shaft of the motor 94 and an air stream surrounds the propeller. The situation appears during the flight of the airplane. [00164] Fig.21 illustrates a motor controller 93 with a DC voltage Vdc at the input side and a motor 94 connected at the output side. The motor controller 93 comprises multiple switches 931-933 configured to interrupt or arrange the connection between electrical circuits (not shown) comprised in the motor controller 93. The electrical circuits can comprise power semiconductors for converting the DC voltage VDC at the input side into alternating quantities, such as alternating voltages and/or currents. [00165] The switches 931-933 may be high power contactors. Other forms of switches, such as semiconductor switches, might be used instead. [00166] A control circuit (not shown) is connected to the switches 931-933 and configured to control the said switches from a non-conductive state into a conductive state and vice versa. The control circuit may include digital components, including for example a processor, a FPGA circuit, and/or any combination of digital and/or analogue components for controlling the switches. Also, in this example, passes the airstream the propeller. [00167] A speed sensor in the form of a rotary encoder is mechanically connected to the motor shaft, as in the examples illustrated in Figs.20 and 21. The angular position of the rotor of the motor 94 can be determined with the use of the said rotary encoder. In addition, the speed of the rotor (angular frequency) can be determined. The rotary encoder is electrically connected to the control circuit of Fig.21. [00168] Fig.22 illustrates a simplified control diagram for a motor controller 93 connected to the motor 94 of Figs.21 and 22. H55-24-PCT [00169] The motor controller includes an inverter circuit 945 configured to convert the DC voltage Vdc at the input side into three-phase alternating currents and voltages at the output side. Three phase lines interconnect the inverter circuit 945 of the motor controller 93 and the motor 94. Each of the three phase lines includes a switch 931-933 connected in series with each phase line. Each switch 931-933 is connected to the control circuit 950 and can be controlled by the control circuit 950 between a conductive and non-conductive state and vice versa. [00170] In this example, each phase line further includes a plurality of current sensors 961-963 between the inverter circuit 945 and the switches 931-933, for measuring an electrical current flowing in each of the phase lines during the operation of motor 94. Each current sensor 961-963 is connected to the control circuit 950. The current sensors 961-963 may be hall effect current transducers. Other types of current sensors, such as current transformers, Rogowski coils, or even shunt resistors, might be used instead. The voltage of each phase line is also measured with the use of a plurality of voltage sensors 964-966, such as voltage transducers. [00171] The motor 94 includes the rotary encoder 960 mechanically connected to the shaft of the motor for determining the rotor speed ωM and the rotor position Θ. The rotary encoder 960 is electrically connected to the control circuit 950. The rotor speed ωM and the rotor position Θ can be alternatively calculated and/or estimated by the control circuit 950 using a position and speed estimator 943. [00172] The position and speed estimator 943 calculates and/or estimates the rotor speed ωM^* and the rotor position Θ^* based on measured phase currents or phase voltages and a machine model known from the prior art. The position and speed estimator 943 is essential for the operation of the motor controller 93 and motor 94 in case of a failure of the rotary encoder 960 and will be explained in greater detail in the course of this example. H55-24-PCT [00173] Three control modes are considered. In the sensor-based control mode the motor controller 93 controls the motor 94 with the use of a sensor-based vector control method utilizing the sensor signals provided by the rotary encoder 960. In case of a loss of the sensor signal or a defect of the rotary encoder 960, a sensorless control mode needs to be considered, wherein the position of the rotor needs to be estimated from electric driving signals. The same control mode might be used during the start-up of the motor controller 93, in particular when the aircraft is grounded, and the exact rotor position is not known. The initialization control mode is considered when the sensor signal of the rotary encoder 960 is permanently lost or disregarded for other purposes, such as a defect of the rotary encoder 960. In this mode, the motor controller 93 controls the motor 94 permanently with the sensorless vector control method. [00174] The control diagram illustrates key elements for a field-oriented control of the motor 94 only. A field-oriented controller (FOC) may be implemented, for example, as a software program contained in a memory of the control circuit 950, as a FPGA, or with other means. As said, the FOC can be used for all control modes in different variations. [00175] In case of software implementation, a software program is executed by a processor as part of the control circuit 950 during operations of the motor controller 93 and configured to control the electrical energy provided to the motor 94 in a control loop. [00176] The FOC generates a three-phase voltage as a vector vS to control the three-phase stator current of the motor 94. The stator currents comprise two orthogonal components that can be represented with a vector. One component defines the magnetic flux of the motor 94, the other component of the vector the torque. By transforming the AC currents into rotational vectors using transforms, the flux and the torque components become time-invariant and thus allow the control with conventional techniques such as PI controllers, as with a standard DC motor. The term “controller”, such as flux controller, relates in the H55-24-PCT following to a software or hardware module, dependent on how the FOC is implemented. [00177] In the following the term “value” represents electrical or physical quantities determined by measurement or set by the motor controller 93, whereas the term “variable” is a result of a calculation or transformation of a value represented in the processor. [00178] During the sensor-based control mode, all three-phase stator currents of the motor 94 are determined by the current sensors 961-963 connected to the control circuit 950. These measurements provide values iU, iV and iW. In the prior art, it is often found that only two phase currents are measured and a third phase current is calculated based to the relationship iU+iV+iW = 0. [00179] The three-phase stator currents of the motor 94 are converted to a two-axis coordinate system using the Clarke transformation 940. This conversion provides the variables iα and iβ from the measured AC stator currents iU - iW. The variables iα and iβ are time-varying quadrature current values as viewed from the perspective of the stator. The rotor position Θ is directly measured by the encoder 960 or derived by integrating the speed determined by the encoder 960. [00180] The two-axis coordinate system is rotated to align with the rotor flux using a transformation angle calculated at an initial or previous iteration of the control loop. The Park transformation 941, using the rotor position Θ, provides the id and iq variables derived from the variables iα and iβ. The id and iq variables are the quadrature currents transformed to the rotating coordinate system. For steady-state conditions, id and iq are constant. [00181] A speed set point ωRef corresponding to the rotor speed is set and an error signal is formed using the speed set point ωRef and the determined the rotor speed ωM^*. The velocity controller 948 is provided as a PI- H55-24-PCT controller and regulates its output, being the iq variable according to the error signal. [00182] The speed of the rotor cannot be increased above the rated speed of the motor 94, due to saturation of the ferromagnetic part of the motor with the magnetic flux generated by the rotation of the rotor. However, with the use of the field weakening controller 949 the torque of the motor 94 can be exhaustively utilized in all operational ranges. The idea of a field weakening controller 949 is to lower the resulting d-flux component (variable id, rotor magnetizing flux) by reducing the effect of the flux of the rotor. [00183] Therefore, the field weakening controller 949 controls the id variable. By setting the speed set point ωRef above the rated speed of the motor 94, the velocity controller 948 increases the q-component (variable iq, torque output) of the flux, whereas the field weakening controller reduces the d-flux component at the same time. [00184] Further error signals are formed using the variables id, iq and corresponding set points id^*, iq^*. The set point id^* controls the rotor magnetizing flux and the set point iq^*controls the torque output of the motor 94. The error signals are inputted to flux controller 947 and into the torque controller 946, wherein each is configured as PI controller. Other controller types, such as a bang-bang controller might be used instead. [00185] The output of the flux controller 947 and the torque controller 946 provide the variables vd and vq, representing a voltage vector with two voltage vector components that will be set to the motor 94. The two voltage component vectors may be represented in the rotating d-q axis. [00186] A new transformation angle is calculated in a subsequent iteration of the control loop, where the variables vα, vβ, iα and iβ are considered as inputs. The new transformation angle guides the FOC as to where to place the next voltage vector vS. H55-24-PCT [00187] The variables vd and vq provided by the flux- and torque controller 947, 946 are rotated back to the stationary reference frame using the new transformation angle. The inverse Park transformation 942 provides the subsequent quadrature voltage values vα and vβ under consideration of the current rotor position Θ. [00188] The subsequent quadrature voltage variables vα and vβ are transformed back to three-phase voltage values using an inverse Clarke transformation 942 in the Pulse-Width Modulation (PWM) Modulator 944. New PWM duty cycle values vUC-vWC are calculated in the PWM modulator based on the transformed three-phase voltage values for signalling the inverter circuit 945. [00189] The inverter circuit 945 sets its pulse pattern according to the provided duty cycle values vUC-vWC. [00190] This process and the corresponding control loop are executed periodically, as long as the motor controller 93 provides the motor 94 with electrical energy. [00191] In the sensorless control mode, the speed signal provided by the rotary encoder 960 is not available and/or is disregarded for other purposes. The loss of the sensor signal can occur during the sensor-based control mode, in particular when the motor controller 93 controls the motor 94 with the use of the sensor-based vector control method. [00192] In this case, the control circuit 950 needs to transition into the sensorless control mode. In this mode the motor controller 93 can control the motor 94 using a V/f or I/f open loop control. This can be achieved by using the control diagram as illustrated but disabling the feedback loop (the lower part in the diagram, including the Clarke transformation 940, the Park transformation 941, and the position and speed estimator 943) and thereby directly influencing the variables vd and vq (V/f control) or id and iq (I/f control) and finally the voltage vector vS. H55-24-PCT [00193] Controlling the motor with the use of the method is necessary to turn the rotor, for instance, during start (from a standstill to a first rotational speed, sufficient to measure the back-EMF generated by the motor). However, and in case the airplane is cruising, the application of this control method is not required, as the rotor remains turning, due to the airstream passing the propeller. If the V/f or I/f open loop control is used to control the motor 94, then the control is stopped. The switch (as illustrated in Fig 21) is opened and the motor 94 generates a voltage, that can be determined with the voltage sensors 964-966, as the rotor of the motor 94 keeps turning. [00194] This also applies to the situation in which the propeller turns the rotor during flight. The step of interrupting the phase lines can be necessary to ensure that the motor 94 is still turning, and the faulting signal is not caused by the defect of the motor 94. Furthermore, the voltages measured vu-vw can be used to estimate the rotor position Θ* and the rotor speed ωM^* by the position and speed estimator 943 with the use of the measured voltages. [00195] The position and speed estimator 943 is functioning in this case as a simple back-emf observer, only relying on the voltages that are induced by the rotating rotor into the stator coil of the motor 94. For this purpose the measured voltages vu-vw are transformed with the use of the Clarke transformation 940 into variables vα^*, vβ^* in the αβ coordinate system and finally into variables vd^*, vq^* which represents quadrature voltages transformed to the rotating coordinate system (d-q reference frame). The position and speed estimator 943 can then estimate using the variables vd^*, vq^* the rotor position Θ* and the rotor speed ωM^* with methods known from the prior art. [00196] However, interrupting the phase lines is not the only possibility to determine the rotor position in situations of the loss of the signal of the rotary encoder 960. There are further solutions known from the prior art for which it is not necessary to interrupt the phase lines. H55-24-PCT [00197] For instance, the motor controller 93 can comprise a resistor that can be switchable activated (also called braking resistor). When the motor 94 supplies electrical energy, the said energy can be wasted in the said resistor, for the sake of causing determinable currents in the phase lines. [00198] By determining the currents and voltages of the phase lines, the rotor position Θ* and the rotor speed ωM^* can be estimated. The estimation can for example use of the method, such as disclosed in the publication "Zhiqian Chen, M. Tomita, S. Doki, et al. “An extended electromotive force model for sensorless control of interior permanent-magnet synchronous motors”. In: IEEE Transactions on Industrial Electronics 50.2 (Apr.2003), pp. 288–295. https://doi.org/10.1109/TIE.2003.809391/ [00199] Alternatively, the motor controller 93 can feed the energy that is supplied by the motor 94 into the battery and the related currents and voltages measured in the phase lines can also be used to estimate the rotor position Θ* and the rotor speed ωM^*, also with the use of the extended back-EMF observer as disclosed in the said document. [00200] Once the rotor position Θ* and the rotor speed ωM^* is estimated using one of the before outlined possibilities, the initialization control mode is entered and/or used. The motor controller 93 controls the motor 94 with the use of the sensorless bases vector control method, taking benefit of the estimated rotor position Θ* and the rotor speed ωM^* as starting point for the conventional sensorless based motor control. [00201] The example outlined herein before takes benefit of the behavior of the permanent-magnet synchronous motor. However, the control structure is also suitable for other types of electrical machines, such as induction machines or DC machines, but the control structure would require a slight adaptation of the control in the sensorless control mode. H55-24-PCT Additional Features and

[00202] Although examples provided herein may be described in the context of an aircraft, such as an electric or hybrid aircraft, one or more features may further apply to other types of vehicles usable to transport passengers or goods. For example, the one or more futures can be used to enhance construction or operation of automobiles, trucks, boats, submarines, spacecrafts, hovercrafts, or the like. [00203] Many other variations than those described herein will be apparent from this disclosure. For example, depending on the embodiment, certain acts, events, or functions of any of the algorithms described herein can be performed in a different sequence, can be added, merged, or left out altogether (for example, not all described acts or events are necessary for the practice of the algorithms). Moreover, in certain embodiments, acts or events can be performed concurrently, for instance, through multi- threaded processing, interrupt processing, or multiple processors or processor cores or on other parallel architectures, rather than sequentially. In addition, different tasks or processes can be performed by different machines or computing systems that can function together. [00204] Unless otherwise specified, the various illustrative logical blocks, modules, and algorithm steps described herein can be implemented as electronic hardware, computer software, or combinations of both. To clearly illustrate this interchangeability of hardware and software, various illustrative components, blocks, modules, and steps have been described above generally in terms of their functionality. Whether such functionality is implemented as hardware or software depends upon the particular application and design constraints imposed on the overall system. The described functionality can be implemented in varying ways for each particular application, but such implementation decisions should not be interpreted as causing a departure from the scope of the disclosure. [00205] Unless otherwise specified, the various illustrative logical blocks and modules described in connection with the embodiments disclosed H55-24-PCT herein can be implemented or performed by a machine, a microprocessor, a state machine, a digital signal processor (DSP), an application specific integrated circuit (ASIC), a FPGA, or other programmable logic device, discrete gate or transistor logic, discrete hardware components, or any combination thereof designed to perform the functions described herein. A hardware processor can include electrical circuitry or digital logic circuitry configured to process computer-executable instructions. In another embodiment, a processor includes an FPGA or other programmable device that performs logic operations without processing computer-executable instructions. A processor can also be implemented as a combination of computing devices, e.g., a combination of a DSP and a microprocessor, a plurality of microprocessors, one or more microprocessors in conjunction with a DSP core, or any other such configuration. A computing environment can include any type of computer system, including, but not limited to, a computer system based on a microprocessor, a mainframe computer, a digital signal processor, a portable computing device, a device controller, or a computational engine within an appliance, to name a few. [00206] Unless otherwise specified, the steps of a method, process, or algorithm described in connection with the embodiments disclosed herein can be embodied directly in hardware, in a software module stored in one or more memory devices and executed by one or more processors, or in a combination of the two. A software module can reside in RAM memory, flash memory, ROM memory, EPROM memory, EEPROM memory, registers, hard disk, a removable disk, a CD-ROM, or any other form of non-transitory computer-readable storage medium, media, or physical computer storage known in the art. An example storage medium can be coupled to the processor such that the processor can read information from, and write information to, the storage medium. In the alternative, the storage medium can be integral to the processor. The storage medium can be volatile or nonvolatile. The processor and the storage medium can reside in an ASIC. [00207] Conditional language used herein, such as, among others," can,” "might,” "may,” "e.g.," and the like, unless specifically stated otherwise, or H55-24-PCT otherwise understood within the context as used, is generally intended to convey that certain embodiments include, while other embodiments do not include, certain features, elements or states. Thus, such conditional language is not generally intended to imply that features, elements or states are in any way required for one or more embodiments or that one or more embodiments necessarily include logic for deciding, with or without author input or prompting, whether these features, elements or states are included or are to be performed in any particular embodiment. The terms "comprising,” "including,” "having," and the like are synonymous and are used inclusively, in an open-ended fashion, and do not exclude additional elements, features, acts, operations, and so forth. Also, the term "or" is used in its inclusive sense (and not in its exclusive sense) so that when used, for example, to connect a list of elements, the term "or" means one, some, or all of the elements in the list. Further, the term "each," as used herein, in addition to having its ordinary meaning, can mean any subset of a set of elements to which the term "each" is applied. Reference symbols in the figures 91 First battery pack 92 Second battery pack 93 Motor controller 94 Motor 931, 932, 933 Switch 940 Clarke transformation 941 Park transformation 942 Inverse Park transformation 943 Rotor position and speed estimator 944 PWM modulator 945 Inverter circuit 946 Torque controller 947 Flux controller 948 Velocity controller 949 Field weakening controller 950 Control circuit H55-24-PCT 960 Resolver 961, 962, 963 Current sensors 964, 964, 966 Voltage sensors id Rotor magnetizing flux (variable or value) iq Rotor toque, torque output (variable or value) iU, iV, iW Current phase U, V, W vU, vV, vW Phase voltages U, V, W iα, iβ Time-varying quadrature current (variable or value) in the αβ-coordinate system vα, vβ Time-varying quadrature voltages (variable or value) in the αβ-coordinate system iS Stator current vector (absolute value) vd, vq Components of voltage vector vS (variable or value) vd*, vq* Quadrature voltage (variable or value) in the dq-coordinate system vα, vβ Quadrature voltage (variable or value) in the the αβ-coordinate system VDC,vdc DC voltage, DC link voltage vS Stator voltage vector (absolute value) Θ Rotor position (measured) Θ* Rotor position (estimated) ω, ωM, ωM^* Rotor speed, actual rotor speed, estimated rotor speed ωMax Maximum rotor speed ωRef, ωr Speed set point, Rated speed H55-24-PCT

Claims

Claims 1. Motor controller (93) for an electrically or hybrid driven aircraft (100), comprising: - an input end configured to receive electrical energy from an electrical source (91, 92); - an output end configured to supply electrical energy to and receive electrical energy from a motor (94); - a signal input configured to receive signals from a speed or position sensor (960), wherein the motor controller (93) is arranged to provide electrical energy at the output end in the form of a set of driving signals with variable voltage and variable frequency with the use of different control modes for controlling the set of driving signals, wherein in a sensor-based control mode the motor controller (93) is configured to control the set of driving signals at the output end based on a first control scheme, wherein the first control scheme is configured to utilize the signals at the signal input and in a sensorless control mode the motor controller (93) is configured to control the set of driving signals with the use of a second control scheme where the signals at the signal input are disregarded. 2. The motor controller (93) of claim 1, configured to determine or estimate a rotor position of the rotor by determining a stator current or voltage when the rotor rotates and the motor generates electrical energy, and is further arranged to initialize the sensorless control mode based on the determined or estimated rotor position. 3. The motor controller (93) of claim 1 or 2, wherein the second control scheme is based on a V/f- or I/f-control, whereby V denotes the voltage, I denotes the current, and f denotes the frequency of the set of driving signals at the output end. H55-24-PCT 4. The motor controller (93) of claim 1, wherein the first control scheme is based on sensor-based vector control and the second control scheme is based on sensorless vector control. 5. The motor controller (93) of one of the claims 1 to 4, configured with an initialization control mode, in which the motor controller (93) is arranged to receive electrical energy at the output end and configured to estimate a virtual rotor position from the electrical energy received. 6. The motor controller (93) of any one of the claims 1 to 5, configured to detect a failure at the signal input, wherein the motor controller (93) is arranged to transition from the sensor control mode into the sensorless control mode upon the detection of the failure. 7. The motor controller (93) of claim 6, configured to detect a failure at the signal input, wherein the motor controller (93) is arranged to transition from the sensor mode, then into the initialization mode and subsequently into the sensorless control mode upon the detection of the failure. 8. The motor controller (93) of claim 7, configured to utilize the determined or estimated rotor position for the second control scheme while transitioning from the sensor-based or from the initialization control mode into the sensorless control mode. 9. The motor controller (93) of any one of the claims 6 to 8, comprising a control circuit (950) configured to implement and transition between the different control modes, configured to detect the failure at the signal input, and configured to determine or estimate the rotor position. 10.The motor controller (93) of any one of the claims 1 to 9, comprising current sensors and voltage sensors, configured to measure electrical quantities in the set of driving signals and of the electrical energy receivable at the output end. H55-24-PCT 11.The motor controller (93) of any one of the claims 1 to 10, comprising a switching device (931, 932, 933) configured to conduct or insulate the set of driving signals and of the electrical energy receivable at the output end. 12.Propulsion system (900) for an electric or hybrid aircraft (100), comprising: - an electrical source (91, 92); - a first motor (94) having a rotor; - a speed or position sensor (960) coupled to the rotor for measuring a speed and/or a position of the said rotor; - a motor controller (93) of any one of the claims 1 to 11 connected to the electrical source (91, 92) at the input end, connected to the sensor (960) at the signal input, and connected to the first motor (94) at the output end. 13.The propulsion system (900) of claim 12, wherein the first motor (94) is a permanent magnet synchronous motor, the propulsion system comprising a propeller coupled to the rotor. 14.The propulsion system (900) of claim 12 or 13, comprising a second motor (99) mechanically coupled to the rotor of the first motor (94). 15.Method for operating the propulsion system (900) of any one of the claims 12 to 14, comprising the steps of: - optionally controlling the rotor of the first motor (94) from standstill to a operational speed by controlling the set of driving signals of the motor controller (93) using the sensor-based control mode or the sensorless control mode; - detecting a failure at the signal input; - switching the control mode from the sensor-based or sensorless control mode to the initialization control mode, such that a remaining rotation of the rotor of the first motor (94) supplies the motor controller (93) at the output end with electrical energy; - determining or estimating the rotor position of rotor of the first motor (94) from the electrical energy received at the output end; H55-24-PCT - switching the control mode from the initialization control mode to the sensorless control mode; - controlling the set of driving signal at the output end of the motor controller (93) using the determined or estmated rotor position and the sensorless vector control. 16.The method of claim 15, comprising the step of maintaining the operational speed of the rotor by controlling the set of driving signal at the output end of the motor controller (93) using the sensorless vector control. 17.The method of any one of the claims 15 to 16, comprising a step of energizing the second motor (99) during an activation of the initialization control mode, for rotating the rotor of the first motor (94). 18.The method of any one of the claims 15 to 17, comprising a step of changing a blade pitch of the propeller coupled to the rotor during an activation of the initialization control mode . 19.The method of any one of the claims 15 to 18, comprising a step of decoupling the first motor (94) from the motor controller (93) by insulating the set of driving signals with the use of the switching device (931, 932, 933) during an activation of the initialization control mode. 20.The method of any of the claims 15 to 19, wherein the steps are performed in the sequence as listed. 21.Electric or hybrid airplane (100) propelled by the propulsion system of claim 12, wherein the method of any one of the claims 15 to 20 is performed during flight of the airplane (100). H55-24-PCT