An electric vertical take-off and landing aircraft

Abstract

An electric vertical take-off and landing aircraft is described. The electric vertical take-off and landing aircraft comprises a plurality of batteries, a plurality of separate power distribution assemblies, and a plurality of electric propulsion units. Each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies. Each power distribution assembly is connected to multiple electric propulsion units of the plurality of electric propulsion units. Each electric propulsion unit of the plurality of electric propulsion units comprises an electric motor comprising a plurality of windings, a plurality of power control elements, and a controller. Each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings. The controller is configured to receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements, and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

Description

[0001] AN ELECTRIC VERTICAL TAKE-OFF AND LANDING AIRCRAFT

[0002] Field of the Invention

[0003] The present invention relates to an Electric Vertical Take-Off and Landing (eVTOL) aircraft.

[0004] Background of the Invention

[0005] EVTOL aircraft are aircraft which take off vertically, transition to conventional, horizontal flight, and then land vertically. There is a general desire to improve the fault tolerance, battery management, and / or reduce the weight of eVTOL aircraft.

[0006] Summary of the Invention

[0007] According to a first aspect of the present invention there is provided an electric vertical takeoff and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, and a plurality of electric propulsion units, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements; and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

[0008] By performing the control within each Electric Propulsion Unit (EPU), the responsiveness of the controller may be improved when compared with sending the measurements to a central flight control computer, the central flight computer processing the measurements (among many other tasks that it is performing) and then the central flight control computer sending commands to control the power control elements. For example, the inventors have identified that response times can be reduced significantly (e.g. from 10ms to 1ms) when performing the control within the EPU and avoiding the processing overhead and communications latency of using the central flight control computer. This may be particularly beneficial in an eVTOL aircraft, where the controller may perform hundreds of control iterations a second in order to maintain the stability of the eVTOL aircraft.

[0009] Connecting each battery to a separate power distribution assembly may provide redundancy in the power distribution system of the eVTOL aircraft such that the failure of a single power distribution assembly may not result in the loss of ability to distribute power from all the batteries, as may occur if the batteries were connected to a single, common power distribution assembly (for example, a common busbar).

[0010] Connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery may improve the fault tolerance of the aircraft. For example, in the event of a power distribution assembly or battery failure, the winding not connected to the failed power distribution assembly or battery may continue to drive the electric motor. Additionally, in the event of an EPU failure, each battery may continue to provide power to a non-failed EPU. In contrast, if each battery were connected to only a single EPU, the mass of the batteries may need to be greater to accommodate for the lack of utilisation of the charge within the battery connected to the failed EPU.

[0011] By controlling the level of power delivered to the repective windings to reducing the difference in power delivery properties (such as discharge rate, State of Charge (SoC), or temperature), the controller may be able to reduce differences in power delivery properties between the batteries which may occur during a flight. For example, in the event that the discharge rate of one of the batteries connected to one of the windings is lower than the discharge rate of the other battery connected to the other winding of the plurality of windings (for example, if one of the multiple EPUs the battery is powering fails), the controller may instruct the plurality of power control elements to draw less power from the battery with the higher discharge rate. This may reduce the difference between the discharge rates and thereby reduce the difference in the SoCs between the batteries. Reducing the difference in SoC between the batteries may enable the overall weight of batteries to be reduced because additional battery capacity may not be required to compensate for the under utilisation of some of the batteries.

[0012] Additionally, reducing the difference in SoC may reduce the likelihood of some of the batteries becoming depleted before others of the batteries. If batteries become depleted prematurely, the aircarft may no longer be able to operate the EPUs connected to the depleted batteries, and mission duration and range may be decreased.

[0013] Optionally, the power delivery property is a discharge rate. By measuring the discharge rate rather than, for example, the SoC, the responsiveness of the controller may be improved. This is because the difference in discharge rates may cause the difference in SoCs. Therefore, a difference in discharge rates may be detectable before a difference in SoCs is detectable and thereby the controller may be able to take preventative action sooner. Improving the responsiveness of the controller may improve the extent to which the SoCs of the batteries can be balanced. Additionally, the discharge rate may be measurable close to the power control elements, rather than close to the battery (which may be located in a different part of the aircraft to the EPUs). This may reduce the amount of cabling required to connect the sensors to the controller and thereby the weight and manufacturing complexity of the aircraft.

[0014] Optionally, each measurement of the plurality of measurements is a measurement of a voltage of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements. Measuring the voltage can be achieved in an efficient and stable manner and may be simpler than measuring, for example, the current between the respective battery and the respective power distribution assembly. Additionally, the voltage may be measurable close to the power control elements, rather than close to the battery (which may be located in a different part of the aircraft to the EPUs). This may reduce the amount of cabling required to connect the sensors to the controller and thereby the weight and manufacturing complexity of the aircraft.

[0015] Optionally, each power control element of the plurality of power control elements is an inverter; and the electric motor is an alternating current electric motor. Although Direct Current (DC) motors may be employed, Alternating Current (AC) motors tend to have a higher power to weight ratios than DC motors. Thereby, using AC motors may reduce the weight of the aircraft, which may increase the range of the aircraft for a given battery capacity. Optionally, the electric motors are three-phase AC electric motors. Three-phase AC motors tend to have higher power to weight ratios and power factors than single-phase AC motors. As a result, the weight of the electric motor and the batteries may be reduced, which may reduce the weight of the aircraft.

[0016] Optionally, the controller is configured to control the plurality of power control elements according to a regenerative regime when one of the respective batteries of the plurality of batteries connected to each power control element of the plurality of power control elements has a different power delivery property than another of the respective batteries of the plurality of batteries connected to each power control element of the plurality of power control elements; and the regenerative regime comprises the winding of the plurality of windings which is connected to the respective battery with the different power delivery property: regeneratively harvesting energy from the electric motor; and suppling the regeneratively harvested energy to the power distribution assembly connected to the winding which is regeneratively harvesting the energy in order to reduce a difference in power delivery properties indicated by the received plurality of measurements. Thereby, the controller may be able to balance the power delivery property of each battery to a greater extent than if regeneration were not employed. Regeneration may be particularly beneficial in the event of a component failures in which the loss of the component may prevent fully balancing the power delivery properties of the batteries without regeneration. A regenerative regime in this context is to balance power delivery properties and not to recharge a respective battery. In operation, the regenerated energy is coupled to the associated power distribution assembly and is used to increase the power delivered to a neighbouring EPU. In some examples, circuitry (e.g. an arrangement of one or more diodes) may be provided to prevent regenerated energy from reaching an associated battery.

[0017] Optionally, the electric vertical take-off and landing aircraft comprises a flight control computer configured to send a respective power command to the controller of each of the electric propulsion units; and the controller is configured to, in response to receiving the respective power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the respective power command and to reduce a difference in power delivery properties indicated by the received plurality of measurements; and at least two of the respective power commands are different from one another. This may provide greater flexibility in the level of thrust that can be delivered by the different EPUs. For example, this may enable different thrust to be delivered from different EPUs to balance the eVTOL aircraft in order to counteract wind forces and balance the eVTOL aircraft. The inventors have identified that when the EPUs are operating at different powers, and a component failure occurs, regeneration may be particularly beneficial for balancing the power delivery properties of the batteries.

[0018] Optionally, the controller is configured to switch between controlling the plurality of power control elements according to the regenerative regime and controlling the plurality of power control elements according to a proportional regime; and the proportional regime comprises varying the level of power delivered to each winding of the plurality of windings from the respective battery connected to each power control element in proportion to the difference in power delivery properties indicated by the received plurality of measurements. Switching between the two regimes may enable the aircraft to employ the regime which is most able to balance the batteries in a variety of scenarios. For example, the regenerative regime may be more able to balance the batteries after a failure of a components of the eVTOL aircraft has occurred, whilst the proportional regime may be more able to balance the batteries during normal, non-fault, operations.

[0019] Optionally, the controller is configured to switch from controlling the plurality of power control elements according to the proportional regime to controlling the plurality of power control elements according to the regenerative regime based on a maximum power rating of the power control element connected to the other of the respective batteries of the plurality of batteries being greater than a total level of power delivered to the plurality of windings when the controller controls the plurality of power control elements according to the proportional regime. Operating in the regenerative regime may require the level of power drawn from the other of the respective batteries through the power control element connected to the other of the respective batteries to increase to the total level of power delivered to the plurality of windings plus the power which is regeneratively harvested in the proportional regime. This is to prevent a reduction in the power of the EPU and thereby prevent a reduction in thrust from the EPU. Therefore, by switching based on the maximum power rating of the power control element connected to the other winding being greater than the total level of power delivered to the plurality of windings when in the proportional regime, the controller may not switch to the regenerative regime when the maximum power rating would prevent the thrust of the EPU being maintained.

[0020] Optionally, the power delivery property is a discharge rate; the different power delivery property is a higher discharge rate; the controller is configured to switch from controlling the plurality of power control elements according to the proportional regime to controlling the plurality of power control elements according to the regenerative regime based on a discharge rate of the other of the respective batteries of the plurality of batteries connected to each power control element of the plurality of power control elements being less than or equal to a higher discharge rate threshold. At higher discharge rates the likelihood of a battery experiencing a thermal run-away event may increase significantly. Therefore, by switching to the regenerative regime based on the discharge rate of the other of the respective batteries being less than or equal to the higher discharge rate threshold, the likelihood of thermal run aways may be reduced and thereby the safety of the eVTOL aircraft increased.

[0021] Optionally, the electric motor has a maximum power rating; and each power control element of the plurality of power control elements has a maximum power rating of no less than 60% and no greater than 85% of the maximum power rating of the electric motor. Having a higher maximum power rating for each power control element may enable the power control element to deliver a greater proportion of the maximum power of the electric motor in the event of a failure of the other power control element of the plurality or a failure of the battery connected to the other power control element of the plurality. Conversely, having a lower maximum power rating for each power control element may enable the power control elements to be lighter, which may reduce the weight of the aircraft. A maximum power rating of no less than 60% and no greater than 85%, and more specifically no less than 65% and no greater than 75%, may provide a good compromise between these competing factors.

[0022] Optionally, the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and a fore / aft plane which extends parallel to the pitch axis and the yaw axis, perpendicular to the roll axis, and through the centre of gravity; the electric vertical take-off and landing aircraft comprises the flight control computer configured to, in response to detecting a failure of one of the electric propulsion units of the plurality of electric propulsion units, send a power command to the controller of at least one of the electric propulsion units; and the controller is configured to, in response to receiving the power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command and to reduce a difference in power delivery properties indicated by the received plurality of measurements. Thereby, the aircraft may be able to compensate for the failure of an EPU and the associated loss of thrust by varying the thrust of the remaining EPUs.

[0023] Optionally, the flight control computer is configured to: in response to detecting a failure of an electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the left / right plane, send the power command to another electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane, and wherein the power command is a command to increase the power of the electric motor; and / or in response to detecting a failure of a further electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the fore / aft plane, send the power command to a different electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the fore / aft plane, and wherein the power command is a command to increase the power of the electric motor.

[0024] As a result, the flight control computer may compensate for the loss of thrust of the failed EPU and balance the aircraft. However, this may result in the discharge rate of the batteries powering the EPU which receives the power command increasing relative to the other batteries and thereby the SoCs of those batteries depleting at a faster rate. If this were not compensated for, the batteries powering the EPU receiving the power command may become empty during flight and the aircraft no longer able to compensate for the loss of the failed EPU. Alternatively, the weight of the batteries may need to be increased to compensate for this disparity in the discharge rates during an EPU failure event.

[0025] By providing the controller which reduces the difference in power delivery property indicated by the received plurality of measurements, the controller may improve the extent to which the power delivery properties of the batteries can be balanced and thereby reduce differences in SoCs. For example, if the discharge rate of a battery has increased to supply more power to an EPU which has received the power command, the controller of the other EPU connected to the battery may reduce the power draw from the battery and increase the power draw from another battery to which it is connected, thereby balancing the discharge rates, and thereby the SoCs, of the batteries.

[0026] Optionally, the flight control computer is configured to: in response to detecting a failure of an electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane and the first side of the fore / aft plane, send the power command to another electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a second side of the left / right plane and a second side of the fore / aft plane, and wherein the power command is a command to decrease the power of the electric motor. As a result, the flight control computer may compensate for the loss of thrust of the failed EPU and balance the aircraft. As described previously, this may result in a disparity in the power delivery properties of the batteries, which may be balanced by the controller.

[0027] Optionally, for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units are: mounted to the airframe on the same side of the left / right plane and / or the same side of the fore / aft plane; and the only electric propulsion units to which the respective power distribution assembly is connected.

[0028] If a power distribution assembly were connected to EPUs on opposite sides of both planes, in the event of a failure of an EPU and the corresponding reduction in power to the EPU on the opposite side of both planes to balance the aircraft, the battery attached to the power distribution assembly may be isolated between the failed EPU and the powered down EPU. Thereby the battery may be supplying significantly less power that the remaining batteries. This may result in the battery having a higher SoC which may not be utilised by the aircraft during the flight. As the majority of the power distribution assemblies are not connected to multiple EPUs on opposite sides of both planes, the likelihood of a battery becoming isolated and thereby underutilised, may be reduced.

[0029] Optionally, the controller is configured to, in response to a failure of one of the power control elements of the plurality of power control elements, control a non-failed power control element of the plurality of power control elements to increase the level of power that is delivered to the respective winding. As a result, the controller may compensate for the failure of one of the power control elements and balance the aircraft. This may result in the discharge rate of the battery connected to the non-failed power control element increasing relative to the other batteries. As described previously, this disparity may be balanced by the controllers of the other EPUs.

[0030] Optionally, the level of power that is delivered to the respective winding connected to the nonfailed power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the power control element. Increasing to greater than or equal to 90% may enable the EPU comprising the respective winding to deliver the majority of the thrust it was delivering before the failure occurred. As a result, the eVTOL may need to take less extensive power redistribution action, than if the level of power were increased by a lesser amount. Increasing the power by greater than or equal to 90% may be expected to increase the discharge rate of the battery connected to the respective winding and lead to the SoC’s of the batteries being significantly different. However, the inventors have identified that by connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery, the discharge rates of the plurality of batteries can be balanced even when increasing the level of power by greater than or equal to 90%.

[0031] Optionally, the level of power that is delivered to the respective winding connected to the nonfailed power control element is increased to greater than or equal to 95%, 99%, or 99.9% of the total level of power that was delivered to the respective windings immediately before the failure of the power control element.

[0032] Optionally, the controller is configured to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the power control elements, control a non-affected power control element of the plurality of power control elements which is not connected to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding. As a result, the controller may compensate for the failure of one of the batteries or power distribution assemblies and balance the aircraft. This may result in the discharge rate of the battery connected to the non-affected power control element increasing relative to the other batteries. As described previously, this disparity may be balanced by the controllers of the other EPUs.

[0033] Optionally, the level of power that is delivered to the respective winding connected to the nonaffected power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly. Increasing to greater than or equal to 90% may enable the EPU comprising the respective winding to deliver the majority of the thrust it was delivering before the failure occurred. As a result, eVTOL may need to take less extensive power redistribution action, than if the level of power were increased by a lesser amount. Increasing the power by greater than or equal to 90% may be expected to increase the discharge rate of the battery connected to the respective winding and lead to the SoC’s of the batteries being significantly different. However, the inventors have identified that by connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery, the discharge rates of the plurality of batteries can be balanced even when increasing the level of power by greater than or equal to 90%.

[0034] Optionally, the level of power that is delivered to the respective winding connected to the nonaffected power control element is increased to greater than or equal to 95%, 99%, or 99.9% of the total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

[0035] Optionally, the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; and a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units to which a respective power distribution assembly of the majority of the plurality of power distribution assemblies is connected are mounted to the airframe on a same sides of the left / right plane. This may reduce the distances between the batteries and the EPUs and thereby reduce the weight of the cabling and power distribution assemblies. This may also improve the ease of installation of the power distribution assemblies and cabling during manufacture of the aircraft.

[0036] Optionally, the multiple electric propulsion units to which a respective power distribution assembly of the plurality of power distribution assemblies is connected comprise a tilting electric propulsion unit and a fixed electric propulsion unit; the electric vertical take-off and landing aircraft comprises a rotor, a tilting mechanism, and a further rotor; the tilting electric propulsion unit is connected to the rotor; the tilting mechanism is configured to tilt the rotor relative to the airframe between a vertical flight mode in which the rotor delivers thrust which is parallel to the yaw axis and a convention flight mode in which the rotor delivers thrust which is parallel to the roll axis of the aircraft in use; and the fixed electric propulsion unit is connected to the further rotor which is fixed relative to airframe such that, in use, the further rotor delivers thrust which is parallel to the roll axis of the aircraft in use only.

[0037] The thrust demand for conventional flight may be significantly lower than the thrust demand for vertical flight. Therefore, this arrangement may enable both the tilting and fixed EPUs to be used in vertical flight, and then only the tilting EPU to be used in conventional flight. By connecting each power distribution assembly to a tilting EPU and a fixed EPU, the charge of the batteries may be used in both flight modes. In contrast, if some power distribution assemblies were connected to just fixed EPUs, the batteries connected to these power distribution assemblies may only be able to supply power during the vertical flight mode, and therefore may have a higher SoC that batteries which power the tilting EPUs. Therefore, connecting each power distribution assembly to a tilting and a fixed EPU may result in more balanced SoCs between the batteries.

[0038] Optionally, each power distribution assembly of the plurality of power distribution assemblies comprises a busbar. Busbars may have several advantages over other forms of power distribution assemblies (for example, electric cables). For example, busbars may be more easily cooled and thereby may be more appropriate for distributing the high currents used in an eVTOL aircraft. Additionally, in the high current eVTOL application, busbars may be more compact and thereby more easily integrated into an eVTOL’ s airframe than equivalent cables, which may require large bend radii.

[0039] Optionally, (when the controller is not configured to control the plurality of power control elements according to the regenerative regime) each power control element of the plurality of power control elements is unidirectional. A unidirectional power control element may be more electrically efficient than a bidirectional power control element. This may reduce transmission losses which may occur between the batteries and the windings. As a result, the range and / or performance of the eVTOL aircraft may be increased. Additionally, unidirectional power control elements may be simpler and thereby cheaper and less likely to develop faults.

[0040] Optionally, the controller is configured to control the plurality of power control elements, when power is deliverable from the respective batteries of the plurality of batteries to the respective windings, to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings. Power may be deliverable from the respective batteries to the respective windings when there are no faults which inhibit the delivery of power. For example, faults in the electric components of the EPU (such as the respective windings or the power control elements connected to the respective windings), or electric components connected directly to the EPU (such as the power distribution assemblies connected to the respective batteries or the respective batteries). It may be expected that optimum balancing of the power delivery properties of the batteries may be achieved by the level of power delivered to the respective winding being approximately equal to one another when power is deliverable to the respective windings. However, the inventors have identified that during certain circumstances, such as faults in other EPUs, or power distribution assemblies or batteries not directly connected to the EPU, better balancing of the power delivery properties may be achieved by having significantly different levels of power being delivered to the respective windings.

[0041] Optionally, the higher threshold is no less than 90% of the total level of power delivered to the respective windings; and the lower threshold is no greater than 10% of the total level of power delivered to the respective windings. Optionally, the higher threshold is no less than 95%, 99%, or 99.9% of the total level of power delivered to the respective windings and the lower threshold is no greater than 5%, 1%, or 0.1% of the total level of power delivered to the respective windings. The inventors have identified that in certain circumstances, greater differences between the levels of power being delivered to the respective windings may result in better balancing of the power delivery properties of the batteries.

[0042] According to a second aspect of the present invention, there is provided an electric propulsion unit for an electric vertical take-off and landing aircraft, the electric propulsion unit comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

[0043] Optionally, the electric propulsion unit comprises a command connection connectable to a flight control computer of the electric vertical take-off and landing aircraft; and the controller is configured to receive a power command from the command connection; and control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command.

[0044] According to a third aspect of the present invention, there is provided a multi-lane inverter apparatus for an electric propulsion unit of an electric vertical take-off and landing aircraft; the multi-lane inverter apparatus comprising: a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of winding connections, each winding connection of the plurality of winding connections is connectable to a respective winding of a plurality of windings of an alternating current electric motor of the electric propulsion unit; a plurality of inverters, each inverter of the plurality of inverters is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding connection of the plurality of winding connections such that each inverter can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery of the plurality of batteries connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; and control the plurality of inverters to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

[0045] Optionally, the multi-lane inverter apparatus comprises a command connection connectable to a flight control computer of the electric vertical take-off and landing aircraft; and the controller is configured to receive a power command from the command connection; and control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command.

[0046] According to a fourth aspect of the present invention, there is provided an electric vertical takeoff and landing aircraft comprising: an airframe; a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electric propulsion units, a centre of gravity; a pitch axis, a roll axis, and a yaw axis; a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; a fore / aft plane which extends parallel to the pitch axis and the yaw axis, perpendicular to the roll axis, and through the centre of gravity; a rotor; a tilting mechanism, and a further rotor, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly of the plurality of power distribution assemblies is connected to a respective tilting electric propulsion unit and a respective fixed electric propulsion unit of the plurality of electric propulsion units; the tilting electric propulsion is connected to the rotor; the tilting mechanism is configured to tilt the rotor relative to the airframe between a vertical flight mode in which the rotor delivers thrust which is parallel to the yaw axis and a convention flight mode in which the rotor delivers thrust which is parallel to the roll axis of the aircraft in use; the fixed electric propulsion is connected to the further rotor which is fixed relative to airframe such that, in use, the further rotor delivers thrust which is parallel to the roll axis of the aircraft in use only; the respective tilting electric propulsion unit and the respective fixed electric propulsion unit are the only electric propulsion units to which the respective power distribution assembly is connected; and for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the respective tilting electric propulsion unit and the respective fixed electric propulsion unit are mounted to the airframe on the same side of the left / right plane and / or the same side of the fore / aft plane. Optionally, each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings.

[0047] Optionally, each electric propulsion unit of the plurality of electric propulsion units comprises a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements; and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

[0048] According to a fifth aspect of the present invention there is provided an electric vertical takeoff and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electric propulsion units, and a control system, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to a plurality of electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings; and the control system is configured to: for one of the electric propulsion units of the plurality of electric propulsion units, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the plurality of power control elements, control a non-affected power control element of the plurality of power control elements which is not connected to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly; and for the other electric propulsion units of the plurality of electric propulsion units: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a property of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements; and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in the properties indicated by the received plurality of measurements.

[0049] According to a sixth aspect of the present invention there is provided an electric propulsion unit for an electric vertical take-off and landing aircraft, the electric propulsion unit comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the plurality of power control elements in use, control a non-affected power control element of the plurality of power control elements which is not connected in use to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

[0050] According to a seventh aspect of the present invention, there is provided a multi-lane inverter apparatus for an electric propulsion unit of an electric vertical take-off and landing aircraft; the multi-lane inverter apparatus comprising: a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of winding connections, each winding connection of the plurality of winding connections is connectable to a respective winding of a plurality of windings of an alternating current electric motor of the electric propulsion unit; a plurality of inverters, each inverter of the plurality of inverters is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding connection of the plurality of winding connections such that each inverter can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the plurality of inverters in use, control a non-affected inverter of the plurality of inverters which is not connected in use to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

[0051] According to an eight aspect of the present invention there is provided an electric vertical takeoff and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electric propulsion units, and a control system wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; a plurality of power control elements, each power control element of the plurality of power control elements is unidirectional and connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings; and the control system is configured, for at least one electric propulsion unit of the plurality of electric propulsion units, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements; and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

[0052] According to a ninth aspect of the present invention, there is provided an electric propulsion unit for an electric vertical take-off and landing aircraft, the electric propulsion unit comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is unidirectional and connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

[0053] According to a tenth aspect of the present invention, there is provided a multi-lane inverter apparatus for an electric propulsion unit of an electric vertical take-off and landing aircraft; the multi-lane inverter apparatus comprising: a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of winding connections, each winding connection of the plurality of winding connections is connectable to a respective winding of a plurality of windings of an alternating current electric motor of the electric propulsion unit; a plurality of inverters, each inverter of the plurality of inverters is unidirectional and connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding connection of the plurality of winding connections such that each inverter can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery of the plurality of batteries connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; and control the plurality of inverters to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

[0054] According to a eleventh aspect of the present invention there is provided an electric vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electric propulsion units (EPUs), and a control system, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple EPUs of the plurality of EPUs; each EPU of the plurality of EPUs comprises: an electric motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding, wherein the control system is configured, for an EPU of the plurality of EPUs, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery connected to a respective power control element of the EPU; and control a power control element of the EPU according to a regenerative regime when the respective battery to which the power control element is connected has a higher discharge rate than the, or each, other respective battery connected to the, or each, other power control element of the EPU, wherein the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

[0055] As a result, the control system may be able to balance the discharge rate of each battery to a greater extent than if regeneration were not employed. Regeneration may be particularly beneficial in the event of a component failures in which the loss of the component may prevent fully balancing the discharge rates of the batteries without regeneration. A regenerative regime in this context is to balance discharge rates and not to recharge a respective battery. In operation, the regenerated energy is coupled to the associated power distribution assembly and is used to increase the power delivered to a neighbouring EPU. In some examples, circuitry (e.g. an arrangement of one or more diodes) may be provided to prevent regenerated energy from reaching an associated battery.

[0056] By using the regenerative regime to reduce the difference in discharge rates , the control system may be able to reduce differences in the state of charge (SOC) of the batteries which may occur during a flight. For example, in the event that the discharge rate of one of the batteries connected to one of the windings is lower than the discharge rate of the other battery connected to the other winding of the plurality of windings , the control system may instruct the plurality of power control elements to supply the regeneratively harvested energy to the power distribution assembly connected to the battery with the higher discharge rate. This may reduce the difference between the discharge rates and thereby reduce the difference in the SoCs between the batteries. Reducing the difference in SoC between the batteries may enable the overall weight of batteries to be reduced because additional battery capacity may not be required to compensate for the under utilisation of some of the batteries.

[0057] Additionally, reducing the difference in SoC may reduce the likelihood of some of the batteries becoming depleted before others of the batteries. If batteries become depleted prematurely, the aircarft may no longer be able to operate the EPUs connected to the depleted batteries, and mission duration and range may be decreased.

[0058] Connecting each battery to a separate power distribution assembly may provide redundancy in the power distribution system of the eVTOL aircraft such that the failure of a single power distribution assembly may not result in the loss of ability to distribute power from all the batteries, as may occur if the batteries were connected to a single, common power distribution assembly (for example, a common busbar).

[0059] Connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery may improve the fault tolerance of the aircraft. For example, in the event of a power distribution assembly or battery failure, the winding not connected to the failed power distribution assembly or battery may continue to drive the motor. Additionally, in the event of an EPU failure, each battery may continue to provide power to a none failed EPU. In contrast, if each battery were connected to only a single EPU, the mass of the batteries may need to be greater to accommodate for the lack of utilisation of the charge within the battery connected to the failed EPU.

[0060] By measuring the discharge rate rather than, for example, the SoC, the responsiveness of the control system may be improved. This is because the difference in discharge rates may cause the difference in SoCs. Therefore, a difference in discharge rates may be detectable before a difference in SoCs is detectable and thereby the control system may be able to take preventative action sooner. Improving the responsiveness of the control system may improve the extent to which the SoCs of the batteries can be balanced. Additionally, the discharge rate may be measurable close to the power control elements, rather than close to the battery (which may be located in a different part of the aircraft to the EPUs). This may reduce the amount of cabling required to connect the sensors to the control system and thereby the weight and manufacturing complexity of the aircraft.

[0061] Optionally, the control system is configured, for the EPU, to switch between controlling the power control element according to the regenerative regime and controlling the power control element according to a proportional regime; and the proportional regime comprises the power control element varying the level of power delivered to the respective winding connected to the power control element from the respective battery connected to the power control element in proportion to the difference in discharge rates indicated by the received plurality of measurements.. Switching between the two regimes may enable the aircraft to employ the regime which is most able to balance the batteries in a variety of scenarios. For example, the regenerative regime may be more able to balance the batteries after a failure of a components of the eVTOL aircraft has occurred, whilst the proportional regime may be more able to balance the batteries during normal, non-fault, operations. Optionally, the control system is configured, for the EPU, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based on a maximum power rating of another power control element of the EPU being greater than a total level of power delivered to the plurality of windings when the control system controls the power control element according to the proportional regime. Operating in the regenerative regime may require the level of power drawn from the other of the respective batteries through the power control element connected to the other of the respective batteries to increase to the total level of power delivered to the plurality of windings plus the power which is regeneratively harvested in the proportional regime. This is to prevent a reduction in the power of the EPU and thereby prevent a reduction in thrust from the EPU. Therefore, by switching based on the maximum power rating of the power control element connected to the other winding being greater than the total level of power delivered to the plurality of windings when in the proportional regime, the control system may not switch to the regenerative regime when the maximum power rating would prevent the thrust of the EPU being maintained.

[0062] Optionally, the control system is configured, for the EPU, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based a discharge rate of the respective battery connected to another power control element of the EPU being less than or equal to a higher discharge rate threshold. At higher discharge rates the likelihood of a battery experiencing a thermal run-away event may increase significantly. Therefore, by switching to the regenerative regime based on the discharge rate of the other of the respective batteries being less than or equal to the higher discharge rate threshold, the likelihood of thermal run aways may be reduced and thereby the safety of the eVTOL aircraft increased.

[0063] Optionally, the electrical vertical take-off and landing aircraft comprises a flight control computer configured to send respective power commands for at least two EPUs of the plurality of EPUs to the control system; the control system is configured, for each EPU of the at least two EPUs to, in response to receiving the respective power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the respective power command and to reduce a difference in discharge rates indicated by the received plurality of measurements; and at least two of the respective power commands are different from one another. This may provide greater flexibility in the level of thrust that can be delivered by the different EPUs. For example, this may enable different thrust to be delivered from different EPUs to balance the eVTOL aircraft in order to counteract wind forces and balance the eVTOL aircraft. The inventor’s have identified that when the EPUs are operating at different powers, and a component failure occurs, regeneration is particularly beneficial for balancing the discharge rates of the batteries.

[0064] Optionally, each measurement of the plurality of measurements is a measurement of a voltage of a respective battery connected to a respective power control element of the EPU. Measuring the voltage can be achieved in an efficient and stable manner and may be simpler than measuring, for example, the current between the respective battery and the respective power distribution assembly. Additionally, the voltage may be measurable close to the power control elements, rather than close to the battery (which may be located in a different part of the aircraft to the EPUs). This may reduce the amount of cabling required to connect the sensors to the control system and thereby the weight and manufacturing complexity of the aircraft.

[0065] Optionally, each power control element of the plurality of power control elements is an inverter; and the electric motor is an alternating current electric motor. Although Direct Current (DC) motors may be employed, Alternating Current (AC) motors tend to have a higher power to weight ratios than DC motors. Thereby, using AC motors may reduce the weight of the aircraft, which may increase the range of the aircraft for a given battery capacity.

[0066] Optionally, the electric motor is a three-phase alternating current electric motor. Three-phase AC motors tend to have higher power to weight ratios and power factors than single-phase AC motors. As a result, the weight of the motor and the batteries may be reduced, which may reduce the weight of the aircraft.

[0067] Optionally, the electric motor has a maximum power rating; and each power control element of the plurality of power control elements has a maximum power rating of no less than 60% and no greater than 85% of the maximum power rating of the electric motor. Having a higher maximum power rating for each power control element may enable the power control element to deliver a greater proportion of the maximum power of the electric motor in the event of a failure of the other power control element of the plurality or a failure of the battery connected to the other power control element of the plurality. Conversely, having a lower maximum power rating for each power control element may enable the power control elements to be lighter, which may reduce the weight of the aircraft. A maximum power rating of no less than 60% and no greater than 85%, and more specifically no less than 65% and no greater than 75%, may provide a good compromise between these competing factors.

[0068] Optionally, the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and a fore / aft plane which extends parallel to the pitch axis and the yaw axis, perpendicular to the roll axis, and through the centre of gravity; the electric vertical take-off and landing aircraft comprises a flight control computer configured to, in response to detecting a failure of one of the EPUs of the plurality of EPUs, send a power command for the EPU to the control system; and the control system is configured, for the EPU, to, in response to receiving the power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command and to reduce a difference in discharge rate indicated by the received plurality of measurements. Thereby, the aircraft may be able to compensate for the failure of an EPU and the associated loss of thrust by varying the thrust of the remaining EPUs.

[0069] Optionally, the flight control computer is configured such that: in response to detecting a failure of an additional EPU of the plurality of EPUs which is mounted to the airframe on a first side of the left / right plane, the EPU is an alternate EPU of the plurality of EPUs which is mounted to the airframe on the first side of the left / right plane, and wherein the power command is a command to increase the power of the electric motor; and / or in response to detecting a failure of a further EPU of the plurality of EPUs which is mounted to the airframe on a first side of the fore / aft plane, the EPU is a different EPU of the plurality of EPUs which is mounted to the airframe on the first side of the fore / aft plane, and wherein the power command is a command to increase the power of the electric motor.

[0070] As a result, the flight control computer may compensate for the loss of thrust of the failed EPU and balance the aircraft. However, this may result in the discharge rate of the batteries powering the EPU which receives the power command increasing relative to the other batteries and thereby the SoCs of those batteries depleting at a faster rate. If this were not compensated for, the batteries powering the EPU receiving the power command may become empty during flight and the aircraft no longer able to compensate for the loss of the failed EPU. Alternatively, the weight of the batteries may need to be increased to compensate for this disparity in the discharge rates during an EPU failure event.

[0071] By providing the control system which reduces the difference in discharge rate indicated by the received plurality of measurements, the control system may improve the extent to which the discharge rates of the batteries can be balanced and thereby reduce differences in SoCs. For example, if the discharge rate of a battery has increased to supply more power to an EPU which has received the power command, the control system of the other EPU connected to the battery may reduce the power draw from the battery and increase the power draw from another battery to which it is connected, thereby balancing the discharge rates, and thereby the SoCs, of the batteries.

[0072] Optionally, the flight control computer is configured such that: in response to detecting a failure of a failed EPU of the plurality of EPUs which is mounted to the airframe on the first side of the left / right plane and the first side of the fore / aft plane, the EPU is a further EPU of the plurality of EPUs which is mounted to the airframe on a second side of the left / right plane and a second side of the fore / aft plane, and wherein the power command is a command to decrease the power of the electric motor. As a result, the flight control computer may compensate for the loss of thrust of the failed EPU and balance the aircraft. As described previously, this may result in a disparity in the discharge rates of the batteries, which may be balanced by the control system.

[0073] Optionally, for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple EPUs are: mounted to the airframe on the same side of the left / right plane and / or the same side of the fore / aft plane; and the only EPUs to which the respective power distribution assembly is connected.

[0074] If a power distribution assembly were connected to EPUs on opposite sides of both planes, in the event of a failure of an EPU and the corresponding reduction in power to the EPU on the opposite side of both planes to balance the aircraft, the battery attached to the power distribution assembly may be isolated between the failed EPU and the powered down EPU. Thereby the battery may be supplying significantly less power that the remaining batteries. This may result in the battery having a higher SoC which may not be utilised by the aircraft during the flight. As the majority of the power distribution assemblies are not connected to multiple EPUs on opposite sides of both planes, the likelihood of a battery becoming isolated and thereby underutilised, may be reduced.

[0075] Optionally, the control system is configured, for the EPU, to, in response to a failure of one of the power control elements of the plurality of power control elements, control a non-failed power control element of the plurality of power control elements to increase the level of power that is delivered to the respective winding. As a result, the control system may compensate for the failure of one of the power control elements and balance the aircraft. This may result in the discharge rate of the battery connected to the non-failed power control element increasing relative to the other batteries. As described previously, this disparity may be balanced by the control system of the other EPUs.

[0076] Optionally, the level of power that is delivered to the respective winding connected to the nonfailed power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the power control element. Increasing to greater than or equal to 90% may enable the EPU comprising the respective winding to deliver the majority of the thrust it was delivering before the failure occurred. As a result, the eVTOL may need to take less extensive power redistribution action, than if the level of power were increased by a lesser amount. Increasing the power by greater than or equal to 90% may be expected to increase the discharge rate of the battery connected to the respective winding and lead to the SoC’s of the batteries being significantly different. However, the inventors have identified that by connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery, the discharge rates of the plurality of batteries can be balanced even when increasing the level of power by greater than or equal to 90%.

[0077] Optionally, the level of power that is delivered to the respective winding connected to the nonfailed power control element is increased to greater than or equal to 95%, 99%, or 99.9% of the total level of power that was delivered to the respective windings immediately before the failure of the power control element.

[0078] Optionally, the control system is configured, for the EPU, to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the power control elements of the plurality of power control elements, control a non-affected power control element of the plurality of power control elements which is not connected to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding. As a result, the control system may compensate for the failure of one of the batteries or power distribution assemblies and balance the aircraft. This may result in the discharge rate of the battery connected to the non-affected power control element increasing relative to the other batteries. As described previously, this disparity may be balanced by the control system for the other EPUs.

[0079] Optionally, the level of power that is delivered to the respective winding connected to the nonaffected power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly. Increasing to greater than or equal to 90% may enable the EPU comprising the respective winding to deliver the majority of the thrust it was delivering before the failure occurred. As a result, eVTOL may need to take less extensive power redistribution action, than if the level of power were increased by a lesser amount. Increasing the power by greater than or equal to 90% may be expected to increase the discharge rate of the battery connected to the respective winding and lead to the SoC’s of the batteries being significantly different. However, the inventors have identified that by connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery, the discharge rates of the plurality of batteries can be balanced even when increasing the level of power by greater than or equal to 90%.

[0080] Optionally, the level of power that is delivered to the respective winding connected to the nonaffected power control element is increased to greater than or equal to 95%, 99%, or 99.9% of the total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

[0081] Optionally, the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; and a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple EPUs to which a respective power distribution assembly of the majority of the plurality of power distribution assemblies is connected are mounted to the airframe on a same sides of the left / right plane. This may reduce the distances between the batteries and the EPUs and thereby reduce the weight of the cabling and power distribution assemblies. This may also improve the ease of installation of the power distribution assemblies and cabling during manufacture of the aircraft.

[0082] Optionally, the multiple EPUs to which a respective power distribution assembly of the plurality of power distribution assemblies is connected comprise a tilting EPU and a fixed EPU; the electric vertical take-off and landing aircraft comprises a rotor, a tilting mechanism, and a further rotor; the tilting EPU is connected to the rotor; the tilting mechanism is configured to tilt the rotor relative to the airframe between a vertical flight mode in which the rotor delivers thrust which is parallel to the yaw axis and a convention flight mode in which the rotor delivers thrust which is parallel to the roll axis of the aircraft in use; and the fixed EPU is connected to the further rotor which is fixed relative to airframe such that, in use, the further rotor delivers thrust which is parallel to the roll axis of the aircraft in use only.

[0083] The thrust demand for conventional flight may be significantly lower than the thrust demand for vertical flight. Therefore, this arrangement may enable both the tilting and fixed EPUs to be used in vertical flight, and then only the tilting EPU to be used in conventional flight. By connecting each power distribution assembly to a tilting EPU and a fixed EPU, the charge of the batteries may be used in both flight modes. In contrast, if some power distribution assemblies were connected to just fixed EPUs, the batteries connected to these power distribution assemblies may only be able to supply power during the vertical flight mode, and therefore may have a higher SoC that batteries which power the tilting EPUs. Therefore, connecting each power distribution assembly to a tilting and a fixed EPU may result in more balanced SoCs between the batteries.

[0084] Optionally, each power distribution assembly of the plurality of power distribution assemblies comprises a busbar. Busbars may have several advantages over other forms of power distribution assemblies (for example, electrical cables). For example, busbars may be more easily cooled and thereby may be more appropriate for distributing the high currents used in an eVTOL aircraft. Additionally, in the high current eVTOL application, busbars may be more compact and thereby more easily integrated into an eVTOL’ s airframe than equivalent cables, which may require large bend radii. Optionally, the control system is configured, for the EPU, to control the plurality of power control elements of the EPU, when power is deliverable from the respective batteries to the respective windings of the EPU, to vary the level of power delivered to the respective windings in order to reduce a difference in the discharge rates indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

[0085] Power may be deliverable from the respective batteries to the respective windings when there are no faults which inhibit the delivery of power. For example, faults in the electrical components of the EPU (such as the respective windings or the power control elements connected to the respective windings), or electrical components connected directly to the EPU (such as the power distribution assemblies connected to the respective batteries or the respective batteries). It may be expected that optimum balancing of the discharge rates of the batteries may be achieved by the level of power delivered to the respective winding being approximately equal to one another when power is deliverable to the respective windings. However, the inventors have identified that during certain circumstances, such as faults in other EPUs, or power distribution assemblies or batteries not directly connected to the EPU, better balancing of the discharge rates may be achieved by having significantly different levels of power being delivered to the respective windings.

[0086] Optionally, the higher threshold is no less than 90% of the total level of power delivered to the respective windings; and the lower threshold is no greater than 10% of the total level of power delivered to the respective windings. Optionally, the higher threshold is no less than 95%, 99%, or 99.9% of the total level of power delivered to the respective windings and the lower threshold is no greater than 5%, 1%, or 0.1% of the total level of power delivered to the respective windings. The inventor’s have identified that in certain circumstances, greater differences between the levels of power being delivered to the respective windings may result in better balancing of the discharge rates of the batteries.

[0087] Optionally, the control system comprises a plurality of controllers; and each EPU of the plurality of EPUs comprises a respective controller of the plurality of controllers; and each controller is configured, for a respective electrical propulsion unit to: control a power control element of the respective EPU according to a regenerative regime when the respective battery to which the power control element is connected has a higher discharge rate than the, or each, other respective battery connected to the, or each, other power control element of the respective EPU, and the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in discharge rate indicated by the received plurality of measurements. By performing the control within each EPU, the responsiveness of the controller may be improved when compared with sending the measurements to a central flight control computer, the central flight computer processing the measurements (among many other tasks that it is performing) and then the central flight control computer sending commands to control the power control elements. For example, the inventors have identified that response times can be reduced significantly (e.g. from 10ms to 1ms) when performing the control within the EPU and avoiding the processing overhead and communications latency of using the central flight control computer. This may be particularly beneficial in an eVTOL aircraft, where the controller may perform hundreds of control iterations a second in order to maintain the stability of the eVTOL aircraft.

[0088] According to a twelfth aspect of the present invention, there is provided a control system for an electric vertical take-off and landing aircraft, the electrical vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of EPUs, and a control system, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple EPUs of the plurality of EPUs; each EPU of the plurality of EPUs comprises: an electric motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings, wherein the control system is configured, for an EPU of the plurality of EPUs, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery connected to a respective power control element of the plurality of the EPU; and control a power control element of the EPU according to a regenerative regime when the respective battery to which the power control element is connected has a higher discharge rate than the, or each, other respective battery of the plurality of batteries connected to the, or each, other power control element of the electrical propulsion unit, wherein the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in discharge rate indicated by the received plurality of measurements..

[0089] According to a thirteenth aspect of the present invention, there is provided an EPU for an electric vertical take-off and landing aircraft, the EPU comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections in use; and control a power control element of the plurality of power control elements according to a regenerative regime when the respective battery to which the power control element is connected in use has a higher discharge rate than the, or each, other respective battery connected in use to the, or each, other respective power control element of the plurality of power control elements, wherein the regenerative regime comprises the power control element causing the winding that is connected in use to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in use in order to reduce a difference in discharge rate indicated by the received plurality of measurements. Optionally, the EPU comprises a command connection connectable to a flight control computer of the electric vertical take-off and landing aircraft; and the controller is configured to receive a power command from the command connection; and control the power control element to vary the level of power delivered to the respective winding based on the power command.

[0090] According to a fourteenth aspect of the present invention, there is provided a multi-lane inverter apparatus for an EPU of an electric vertical take-off and landing aircraft; the multi-lane inverter apparatus comprising: a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of winding connections, each winding connection of the plurality of winding connections is connectable to a respective winding of a plurality of windings of an alternating current electric motor of the EPU; a plurality of inverters, each inverter of the plurality of inverters is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding connection of the plurality of winding connections such that each inverter can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery of the plurality of batteries connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; and control an inverter of the plurality of inverters according to a regenerative regime when the respective battery to which the inverter is connected in use has a higher discharge rate than the, or each, other respective battery connected in use to the, or each, respective inverter of the plurality of inverters, wherein the regenerative regime comprises the inverter causing the winding that is connected in use to the respective battery with the higher discharge rate to regeneratively harvest energy from the alternating current electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in use in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

[0091] Optionally, the multi-lane inverter apparatus comprises a command connection connectable to a flight control computer of the electric vertical take-off and landing aircraft; and the controller is configured to receive a power command from the command connection; and control the power control element to vary the level of power delivered to the respective winding based on the power command.

[0092] According to a fifteenth aspect of the present invention there is provided an electric vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electrical propulsion units, and a control system, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple electric propulsion units (EPUs) of the plurality of EPUs; each EPU of the plurality of EPUs comprises: an electrical motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings, wherein, the control system is configured, for an EPU of the plurality of EPUs, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power control element of the EPU; and when power is deliverable from the respective batteries to the respective windings of the EPU, control the plurality of power control elements of the EPU to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

[0093] Power may be deliverable from the respective batteries to the respective windings when there are no faults which inhibit the delivery of power. For example, faults in the electrical components of the EPU (such as the respective windings or the power control elements connected to the respective windings), or electrical components connected directly to the EPU (such as the power distribution assemblies connected to the respective batteries or the respective batteries). It may be expected that optimum balancing of the power delivery properties of the batteries may be achieved by the level of power delivered to the respective winding being approximately equal to one another when power is deliverable to the respective windings. However, the inventors have identified that during certain circumstances, such as faults in other EPUs, or power distribution assemblies or batteries not directly connected to the EPU, better balancing of the power delivery properties may be achieved by having significantly different levels of power being delivered to the respective windings.

[0094] Connecting each battery to a separate power distribution assembly may provide redundancy in the power distribution system of the eVTOL aircraft such that the failure of a single power distribution assembly may not result in the loss of ability to distribute power from all the batteries, as may occur if the batteries were connected to a single, common power distribution assembly (for example, a common busbar).

[0095] Connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery may improve the fault tolerance of the aircraft. For example, in the event of a power distribution assembly or battery failure, the winding not connected to the failed power distribution assembly or battery may continue to drive the electric motor. Additionally, in the event of an EPU failure, each battery may continue to provide power to a none failed EPU. In contrast, if each battery were connected to only a single EPU, the mass of the batteries may need to be greater to accommodate for the lack of utilisation of the charge within the battery connected to the failed EPU.

[0096] By controlling the level of power delivered to the repective windings to reducing the difference in power delivery properties (such as discharge rate, State of Charge (SoC), or temperature), the control system may be able to reduce differences in power delivery properties between the batteries which may occur during a flight. For example, in the event that the discharge rate of one of the batteries connected to one of the windings is lower than the discharge rate of the other battery connected to the other winding of the plurality of windings (for example, if one of the multiple EPUs the battery is powering fails), the control system may instruct the plurality of power control elements to draw less power from the battery with the higher discharge rate. This may reduce the difference between the discharge rates and thereby reduce the difference in the SoCs between the batteries. Reducing the difference in SoC between the batteries may enable the overall weight of batteries to be reduced because additional battery capacity may not be required to compensate for the under utilisation of some of the batteries.

[0097] Additionally, reducing the difference in SoC may reduce the likelihood of some of the batteries becoming depleted before others of the batteries. If batteries become depleted prematurely, the aircarft may no longer be able to operate the EPUs connected to the depleted batteries, and mission duration and range may be decreased.

[0098] Optionally, the higher threshold is no less than 90% of the total level of power delivered to the respective windings; and / or the lower threshold is no greater than 10% of the total level of power delivered to the respective windings. Optionally, the higher threshold is no less than 95%, 99%, or 99.9% of the total level of power delivered to the respective windings. In addition, or alternatively, the lower threshold is no greater than 5%, 1%, or 0.1% of the total level of power delivered to the respective windings. The inventors have identified that, in certain circumstances, greater differences between the levels of power being delivered to the respective windings may result in better balancing of the power delivery properties of the batteries.

[0099] Optionally, the power delivery property is a discharge rate. By measuring the discharge rate rather than, for example, the SoC, the responsiveness of the control system may be improved. This is because the difference in discharge rates may cause the difference in SoCs. Therefore, a difference in discharge rates may be detectable before a difference in SoCs is detectable and thereby the control system may be able to take preventative action sooner. Improving the responsiveness of the control system may improve the extent to which the SoCs of the batteries can be balanced. Additionally, the discharge rate may be measurable close to the power control elements, rather than close to the battery (which may be located in a different part of the aircraft to the EPUs). This may reduce the amount of cabling required to connect the sensors to the control system and thereby the weight and manufacturing complexity of the aircraft.

[0100] Optionally, each measurement of the plurality of measurements is a measurement of a voltage of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements. Measuring the voltage can be achieved in an efficient and stable manner and may be simpler than measuring, for example, the current between the respective battery and the respective power distribution assembly. Additionally, the voltage may be measurable close to the power control elements, rather than close to the battery (which may be located in a different part of the aircraft to the EPUs). This may reduce the amount of cabling required to connect the sensors to the control system and thereby the weight and manufacturing complexity of the aircraft.

[0101] Optionally, each power control element of the plurality of power control elements is an inverter; and the electric motor is an alternating current electric motor. Although Direct Current (DC) motors may be employed, Alternating Current (AC) motors tend to have a higher power to weight ratios than DC motors. Thereby, using AC motors may reduce the weight of the aircraft, which may increase the range of the aircraft for a given battery capacity.

[0102] Optionally, the electric motors are three-phase AC electric motors. Three-phase AC motors tend to have higher power to weight ratios and power factors than single-phase AC motors. As a result, the weight of the electric motor and the batteries may be reduced, which may reduce the weight of the aircraft.

[0103] Optionally, the control system is configured, for the EPU, to control a power control elements of the EPU according to a regenerative regime when the respective battery to which the power control element is connected has a different power delivery property the, or each, other respective battery of the plurality of batteries connected to the, or each, other power control element of the EPU; and the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the different power delivery property to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in power delivery properties indicated by the received plurality of measurements. Thereby, the control system may be able to balance the power delivery property of each battery to a greater extent than if regeneration were not employed. Regeneration may be particularly beneficial in the event of a component failures in which the loss of the component may prevent fully balancing the power delivery properties of the batteries without regeneration. A regenerative regime in this context is to balance power delivery properties and not to recharge a respective battery. In operation, the regenerated energy is coupled to the associated power distribution assembly and is used to increase the power delivered to a neighbouring EPU. In some examples, circuitry (e.g. an arrangement of one or more diodes) may be provided to prevent regenerated energy from reaching an associated battery. Optionally, the electrical vertical take-off and landing aircraft comprises a flight control computer configured to send respective power commands for at least two EPUs of the plurality of EPUs to the control system; and the control system is configured, for each EPU of the at least two EPUs of the plurality of EPUs to, in response to receiving the respective power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the respective power command and to reduce a difference in power delivery properties indicated by the received plurality of measurements; and at least two of the respective power commands are different from one another. This may provide greater flexibility in the level of thrust that can be delivered by the different EPUs. For example, this may enable different thrust to be delivered from different EPUs to balance the eVTOL aircraft in order to counteract wind forces and balance the eVTOL aircraft. The inventor’s have identified that when the EPUs are operating at different powers, and a component failure occurs, regeneration is particularly beneficial for balancing the power delivery properties of the batteries.

[0104] Optionally, the control system is configured, for the EPU, to switch between controlling the power control element according to the regenerative regime and controlling the power control elements according to a proportional regime; and the proportional regime comprises the power control element varying the level of power delivered to the winding connected to the power control element from the respective battery connected to the power control element in proportion to the difference in power delivery properties indicated by the received plurality of measurements. Switching between the two regimes may enable the aircraft to employ the regime which is most able to balance the batteries in a variety of scenarios. For example, the regenerative regime may be more able to balance the batteries after a failure of a components of the eVTOL aircraft has occurred, whilst the proportional regime may be more able to balance the batteries during normal, non-fault, operations.

[0105] Optionally, the control system is configured, for the EPU, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based on a maximum power rating of another power control element of the EPU being greater than a total level of power delivered to the plurality of windings when the control system controls the power control element according to the proportional regime. Operating in the regenerative regime may require the level of power drawn from the other of the respective batteries through the power control element connected to the other of the respective batteries to increase to the total level of power delivered to the plurality of windings plus the power which is regeneratively harvested in the proportional regime. This is to prevent a reduction in the power of the EPU and thereby prevent a reduction in thrust from the EPU. Therefore, by switching based on the maximum power rating of the power control element connected to the other winding being greater than the total level of power delivered to the plurality of windings when in the proportional regime, the control system may not switch to the regenerative regime when the maximum power rating would prevent the thrust of the EPU being maintained.

[0106] Optionally, the power delivery property is a discharge rate; the different power delivery property is a higher discharge rate; the control system is configured, for the EPU, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based on a discharge rate of the respective battery connected to another power control element of the EPU being less than or equal to a higher discharge rate threshold. At higher discharge rates the likelihood of a battery experiencing a thermal run-away event may increase significantly. Therefore, by switching to the regenerative regime based on the discharge rate of the other of the respective batteries being less than or equal to the higher discharge rate threshold, the likelihood of thermal run aways may be reduced and thereby the safety of the eVTOL aircraft increased.

[0107] Optionally, the electric motor has a maximum power rating; and each power control element of the plurality of power control elements has a maximum power rating of no less than 60% and no greater than 85% of the maximum power rating of the electric motor. Having a higher maximum power rating for each power control element may enable the power control element to deliver a greater proportion of the maximum power of the electric motor in the event of a failure of the other power control element of the plurality or a failure of the battery connected to the other power control element of the plurality. Conversely, having a lower maximum power rating for each power control element may enable the power control elements to be lighter, which may reduce the weight of the aircraft. A maximum power rating of no less than 60% and no greater than 85%, and more specifically no less than 65% and no greater than 75%, may provide a good compromise between these competing factors. Optionally, the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and a fore / aft plane which extends parallel to the pitch axis and the yaw axis, perpendicular to the roll axis, and through the centre of gravity; the electric vertical take-off and landing aircraft comprises a flight control computer configured to, in response to detecting a failure of one of the EPUs of the plurality of EPUs, send a power command for the EPU to the control system; and the control system is configured, for the EPU, to, in response to receiving the power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command and to reduce a difference in power delivery property indicated by the received plurality of measurements. Thereby, the aircraft may be able to compensate for the failure of an EPU and the associated loss of thrust by varying the thrust of the remaining EPUs.

[0108] Optionally, the flight control computer is configured such that: in response to detecting a failure of an additional EPU of the plurality of EPUs which is mounted to the airframe on a first side of the left / right plane, the EPU is an alternate EPU of the plurality of EPUs which is mounted to the airframe on the first side of the left / right plane, and wherein the power command is a command to increase the power of the electric motor; and / or in response to detecting a failure of a further EPU of the plurality of EPUs which is mounted to the airframe on a first side of the fore / aft plane, the EPU is a different EPU of the plurality of EPUs which is mounted to the airframe on the first side of the fore / aft plane, and wherein the power command is a command to increase the power of the electric motor.

[0109] As a result, the flight control computer may compensate for the loss of thrust of the failed EPU and balance the aircraft. However, this may result in the discharge rate of the batteries powering the EPU which receives the power command increasing relative to the other batteries and thereby the SoCs of those batteries depleting at a faster rate. If this were not compensated for, the batteries powering the EPU receiving the power command may become empty during flight and the aircraft no longer able to compensate for the loss of the failed EPU. Alternatively, the weight of the batteries may need to be increased to compensate for this disparity in the discharge rates during an EPU failure event. By providing the control system which reduces the difference in power delivery property indicated by the received plurality of measurements, the control system may improve the extent to which the power delivery properties of the batteries can be balanced and thereby reduce differences in SoCs. For example, if the discharge rate of a battery has increased to supply more power to an EPU which has received the power command, the control system of the other EPU connected to the battery may reduce the power draw from the battery and increase the power draw from another battery to which it is connected, thereby balancing the discharge rates, and thereby the SoCs, of the batteries.

[0110] Optionally, the flight control computer is configured such that: in response to detecting a failure of a failed EPU of the plurality of EPUs which is mounted to the airframe on the first side of the left / right plane and the first side of the fore / aft plane, the EPU is a further EPU of the plurality of EPUs which is mounted to the airframe on a second side of the left / right plane and a second side of the fore / aft plane, and wherein the power command is a command to decrease the power of the electric motor.

[0111] Optionally, for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple EPUs are: mounted to the airframe on the same side of the left / right plane and / or the same side of the fore / aft plane; and the only EPUs to which the respective power distribution assembly is connected.

[0112] If a power distribution assembly were connected to EPUs on opposite sides of both planes, in the event of a failure of an EPU and the corresponding reduction in power to the EPU on the opposite side of both planes to balance the aircraft, the battery attached to the power distribution assembly may be isolated between the failed EPU and the powered down EPU. Thereby the battery may be supplying significantly less power that the remaining batteries. This may result in the battery having a higher SoC which may not be utilised by the aircraft during the flight. As the majority of the power distribution assemblies are not connected to multiple EPUs on opposite sides of both planes, the likelihood of a battery becoming isolated and thereby underutilised, may be reduced.

[0113] Optionally, the control system is configured, for the EPU, to, in response to a failure of one of the power control elements of the plurality of power control elements, control a non-failed power control element of the plurality of power control elements to increase the level of power that is delivered to the respective winding. As a result, the control system may compensate for the failure of one of the power control elements and balance the aircraft. This may result in the discharge rate of the battery connected to the non-failed power control element increasing relative to the other batteries. As described previously, this disparity may be balanced by the control system of the other EPUs.

[0114] Optionally, the level of power that is delivered to the respective winding connected to the nonfailed power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the power control element. Increasing to greater than or equal to 90% may enable the EPU comprising the respective winding to deliver the majority of the thrust it was delivering before the failure occurred. As a result, the eVTOL may need to take less extensive power redistribution action, than if the level of power were increased by a lesser amount. Increasing the power by greater than or equal to 90% may be expected to increase the discharge rate of the battery connected to the respective winding and lead to the SoC’s of the batteries being significantly different. However, the inventors have identified that by connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery, the discharge rates of the plurality of batteries can be balanced even when increasing the level of power by greater than or equal to 90%.

[0115] Optionally, the level of power that is delivered to the respective winding connected to the nonfailed power control element is increased to greater than or equal to 95%, 99%, or 99.9% of the total level of power that was delivered to the respective windings immediately before the failure of the power control element.

[0116] Optionally, the control system is configured, for the EPU, to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the power control elements of the plurality of power control elements, control a non-affected power control element of the plurality of power control elements which is not connected to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding. As a result, the control system may compensate for the failure of one of the batteries or power distribution assemblies and balance the aircraft. This may result in the discharge rate of the battery connected to the non-affected power control element increasing relative to the other batteries. As described previously, this disparity may be balanced by the control system for the other EPUs.

[0117] Optionally, the level of power that is delivered to the respective winding connected to the nonaffected power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly. As a result, eVTOL may need to take less extensive power redistribution action, than if the level of power were increased by a lesser amount. Increasing the power by greater than or equal to 90% may be expected to increase the discharge rate of the battery connected to the respective winding and lead to the SoC’s of the batteries being significantly different. However, the inventors have identified that by connecting each power distribution assembly to multiple EPUs and connecting each winding of each motor to a respective battery, the discharge rates of the plurality of batteries can be balanced even when increasing the level of power by greater than or equal to 90%.

[0118] Optionally, the level of power that is delivered to the respective winding connected to the nonaffected power control element is increased to greater than or equal to 95%, 99%, or 99.9% of the total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

[0119] Optionally, the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; and a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple EPUs to which a respective power distribution assembly of the majority of the plurality of power distribution assemblies is connected are mounted to the airframe on a same sides of the left / right plane. This may also improve the ease of installation of the power distribution assemblies and cabling during manufacture of the aircraft.

[0120] Optionally, the multiple EPUs to which a respective power distribution assembly of the plurality of power distribution assemblies is connected comprise a tilting EPU and a fixed EPU; the electric vertical take-off and landing aircraft comprises a rotor, a tilting mechanism, and a further rotor; the tilting EPU is connected to the rotor; the tilting mechanism is configured to tilt the rotor relative to the airframe between a vertical flight mode in which the rotor delivers thrust which is parallel to the yaw axis and a convention flight mode in which the rotor delivers thrust which is parallel to the roll axis of the aircraft in use; and the fixed EPU is connected to the further rotor which is fixed relative to airframe such that, in use, the further rotor delivers thrust which is parallel to the roll axis of the aircraft in use only.

[0121] The thrust demand for conventional flight may be significantly lower than the thrust demand for vertical flight. Therefore, this arrangement may enable both the tilting and fixed EPUs to be used in vertical flight, and then only the tilting EPU to be used in conventional flight. By connecting each power distribution assembly to a tilting EPU and a fixed EPU, the charge of the batteries may be used in both flight modes. In contrast, if some power distribution assemblies were connected to just fixed EPUs, the batteries connected to these power distribution assemblies may only be able to supply power during the vertical flight mode, and therefore may have a higher SoC that batteries which power the tilting EPUs. Therefore, connecting each power distribution assembly to a tilting and a fixed EPU may result in more balanced SoCs between the batteries.

[0122] Optionally, each power distribution assembly of the plurality of power distribution assemblies comprises a busbar. Busbars may have several advantages over other forms of power distribution assemblies (for example, electrical cables). For example, busbars may be more easily cooled and thereby may be more appropriate for distributing the high currents used in an eVTOL aircraft. Additionally, in the high current eVTOL application, busbars may be more compact and thereby more easily integrated into an eVTOL’ s airframe than equivalent cables, which may require large bend radii.

[0123] Optionally, (when the control system is not configured to control the plurality of power control elements according to the regenerative regime) each power control element of the plurality of power control elements is unidirectional. A unidirectional power control element may be more electrically efficient than a bidirectional power control element. This may reduce transmission losses which may occur between the batteries and the windings. As a result, the range and / or performance of the eVTOL aircraft may be increased. Additionally, unidirectional power control elements may be simpler and thereby cheaper and less likely to develop faults.

[0124] In some examples, the control system is a central system, for example a part of a flight control computer, that is configured, for each EPU of the plurality of EPUs, to receive a plurality of measurements and control the plurality of power control elements as described above. Alternatively, optionally, the control system comprises a plurality of controllers; each EPU of the plurality of EPUs comprises a respective controller of the plurality of controllers; and each controller is configured, for a respective EPU, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power control element of the EPU; and when power is deliverable from the respective batteries of the plurality of batteries to the respective windings of the EPU, control the plurality of power control elements of the EPU to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings. Each controller thereby performs the ‘control system’ operations described above locally, in a decentralised manner, for its associated EPU. By performing the control within each EPU, the responsiveness of the controller may be improved when compared with sending the measurements to a central flight control computer, the central flight computer processing the measurements (among many other tasks that it is performing) and then the central flight control computer sending commands to control the power control elements. For example, the inventors have identified that response times can be reduced significantly (e.g. from 10ms to 1ms) when performing the control within the EPU and avoiding the processing overhead and communications latency of using the central flight control computer. This may be particularly beneficial in an eVTOL aircraft, where the controller may perform hundreds of control iterations a second in order to maintain the stability of the eVTOL aircraft.

[0125] According to a sixteenth aspect of the present invention, there is provided a control system for an electric vertical take-off and landing aircraft, the electrical vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, and a plurality of EPUs, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple EPUs of the plurality of EPUs; each EPU of the plurality of EPUs comprises: an electric motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding, and wherein the control system configured, for an EPU of the plurality of EPUs, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power control element of the EPU; and when power is deliverable from the respective batteries to the respective windings of the EPU, control the plurality of power control elements of the EPU to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

[0126] According to a seventeenth aspect of the present invention, there is provided an EPU for an electric vertical take-off and landing aircraft, the EPU comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections in use; and when power is deliverable from the respective batteries to the respective windings, control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no greater than a lower threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no less than a higher threshold of the total level of power delivered to the respective windings.

[0127] Optionally, the EPU comprises a command connection connectable to a flight control computer of the electric vertical take-off and landing aircraft; and the controller is configured to receive a power command from the command connection; and control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command.

[0128] According to a eighteenth aspect of the present invention, there is provided a multi-lane inverter apparatus for an EPU of an electric vertical take-off and landing aircraft; the multi-lane inverter apparatus comprising: a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of winding connections, each winding connection of the plurality of winding connections is connectable to a respective winding of a plurality of windings of an alternating current electric motor of the EPU; a plurality of inverters, each inverter of the plurality of inverters is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding connection of the plurality of winding connections such that each inverter can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections in use; and when power is deliverable from the respective batteries to the respective windings, control the plurality of inverters to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no greater than a lower threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no less than a higher threshold of the total level of power delivered to the respective windings. Optionally, the multi-lane inverter apparatus comprises a command connection connectable to a flight control computer of the electric vertical take-off and landing aircraft; and the controller is configured to receive a power command from the command connection; and control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command.

[0129] Optional features of aspects of the present invention may be equally applied to other aspects of the present invention, where appropriate.

[0130] Brief Description of the Drawings

[0131] Figure l is a perspective view of an example eVTOL aircraft;

[0132] Figure 2 is a perspective view of the example eVTOL aircraft showing planes of the example eVTOL aircraft;

[0133] Figure 3 is a schematic view of a battery of the example eVTOL aircraft;

[0134] Figure 4 is a top view of the example eVTOL aircraft;

[0135] Figure 5 is a schematic view of an Electric Propulsion Unit (EPU), a rotor, and a tilting mechanism of the example eVTOL aircraft;

[0136] Figure 6 is a schematic view of electric connections between electric components of the example eVTOL aircraft;

[0137] Figure 7 is a flow chart of steps taken by the example eVTOL aircraft in response to receiving a flight command;

[0138] Figure 8 is a schematic view of power distribution through the example eVTOL aircraft in a first scenario;

[0139] Figure 9a is a schematic view of power distribution through a subsection of the example eVTOL aircraft in a second scenario;

[0140] Figure 9b is a schematic view of power distribution through a subsection of the example eVTOL aircraft in a third scenario;

[0141] Figure 9c is a schematic view of power distribution through a subsection of the example eVTOL aircraft in a fourth scenario;

[0142] Figure 10 is a flow chart of steps taken by the example eVTOL aircraft during take-off and landing of the eVTOL aircraft;

[0143] Figure 11 is a schematic view of power distribution through the example eVTOL aircraft in a fifth scenario; Figure 12 is a schematic view of power distribution through the example eVTOL aircraft in a sixth scenario;

[0144] Figure 13 is a schematic view of power distribution through the example eVTOL aircraft in a seventh scenario;

[0145] Figure 14 is a schematic view of power distribution through the example eVTOL aircraft in a eighth scenario;

[0146] Figure 15 is a schematic view of power distribution through the example eVTOL aircraft in a ninth scenario;

[0147] Figure 16 is a flow chart of a decision loop employed by voltage controllers of the example eVTOL aircraft;

[0148] Figure 17 is a flow chart of an alternative decision loop employed by alternative voltage controllers of the example eVTOL aircraft; and

[0149] Figure 18 is a schematic view of an alternative electric connection arrangement between the electric components of the example eVTOL aircraft.

[0150] Detailed Description of the Invention

[0151] The following examples relate to an eVTOL aircraft having plural EPUs powered by plural batteries. In certain examples, an eVTOL has eight batteries powering eight EPUs, and each EPU has two lanes, each comprising a set of windings connected, via an inverter (of a dual lane inverter for the respective EPU), to a respective separate battery. Each battery is associated with a single, isolated power distribution arrangement (e.g., busbar) and drives one set of windings in each of two separate EPUs. Such an arrangement serves to provide a degree of redundancy against component failure, as will be described. The principles disclosed herein apply equally to other arrangements and numbers of batteries and EPUs, and to EPUs with more than two lanes (e.g. three, four, eight or more lanes).

[0152] In an idealised flight mission scenario, in which each battery has the same state of charge and each EPU is controlled to demand the same power, each inverter would be expected to deliver 50% of the required power to the respective EPU lane; and each battery would therefore be required to deliver 100% of the demand, by way of delivering 50% of the demand to two sets of windings in two respective EPUs. In practical scenarios, however, for example due to performing various manoeuvres, varying weather conditions or even in the event of component (e.g. battery, busbar, invertor or motor) failure, as will be described, different EPUs are controlled to demand different powers. In such circumstances, variation in power demand by the EPUs leads to different levels of battery demand, which, if sustained and / or left unchecked, can lead to batteries discharging at significantly different rates. This is undesirable because, if one or more batteries deplete more quickly than other batteries, the eVTOL may suffer reduced utilisation or range, even if there is significant charge remaining in other batteries. Examples herein aim to address this by providing techniques for equalising power demand on the batteries under different scenarios.

[0153] As will be described in relation to some examples, the dual lane inverters serving respective EPUs are adapted to control a split of power demand from the respective pair of batteries that are powering each EPU in a way that is designed to balance or equalise battery power demand across all batteries in the aircraft. The demand on any battery may be due to other EPUs or any other aircraft devices or systems. In some examples, the inverters are configured and / or are operable in drive mode only (i.e. they are one-way inverters) and, in other examples, the inverters are configured and / or are operable in drive mode and regenerative mode (i.e. they are two-way, or bidirectional, inverters). As will also be described, whether in a one-way or a bidirectional mode of operation, the inverters may provide more than one regime of operation, which may for example be deployed depending on the state of the batteries at a given moment.

[0154] Figure 1 and Figure 2 show an eVTOL aircraft 10. The eVTOL aircraft 10 comprises a centre of gravity 12, a pitch axis 14, a roll axis 16, a yaw axis 18, a left / right plane 20, a fore / aft plane 22, an airframe 24, flight controls 26, a flight control computer 28, eight batteries 30, eight busbars 32 (note that only one busbar is shown in Figure 1 for the sake of clarity), eight EPUs 32, eight rotors 33, and four tilting mechanisms 35.

[0155] The centre of gravity 12, the pitch axis 14, the roll axis 16, and the yaw axis 18, the left / right plane 20, and the fore / aft plane 22 are shown in Figure 2. The left / right plane 20 extends parallel to the roll axis 16, perpendicular to the pitch axis 14 and the yaw axis 18, and through the centre of gravity 12. The fore / aft plane 22 extends parallel to the pitch axis 14 and the yaw axis 18, perpendicular to the roll axis 16, and through the centre of gravity 12.

[0156] The airframe 24 comprises a fuselage 36, a pair of wings 38, and four pylons 40. The fuselage 36 comprises a cockpit 42 and a passenger area 44. The pylons 40 each have a generally cuboid shape. The length of each pylon 40 in a direction parallel to the roll axis 16 is greater than a width of the wings 38 in the direction parallel to the roll axis 16. The wings 38 are fixed to an upper side of the fuselage 36 approximately mid-way along the fuselage 36 and extend perpendicular to the fuselage 36. The pylons 40 are fixed to undersides of the wings 38 and spaced apart from one another along the wings 38 in a direction parallel to the pitch axis 14 such that there are two pylons 40 on either side of the left / right plane 20. A fore pylon portion 90 of each pylon 40 projects from the wings 38 on a fore side 100 of the fore / aft plane 22. An aft pylon portion 92 of each pylon 40 projects from the wings 38 on the aft side 102 of the fore / aft plane 22.

[0157] The flight controls 26 are located within the cockpit 42 and enable a pilot of the eVTOL aircraft 10 to input flight commands, such as to take off.

[0158] The flight control computer 28 is located within the fuselage 36. As will be described below in more detail, the flight control computer 28 receives the flight commands from the flight controls 26 and sends power commands to the EPUs 34.

[0159] The batteries 30 (one of which is shown in isolation in Figure 3) each comprise a plurality of individual battery cells 46 enclosed within a housing 48. The batteries 30 are located within the fuselage 36 and provide power to the electric components of the eVTOL aircraft 10.

[0160] The busbars 32 are electrically isolated from one another (and thereby separate from one another). The busbars 32 are an example of power distribution assemblies. In other examples, other power distribution assemblies (for example, electric cables) may be used.

[0161] Referring now to Figure 4, the eight EPUs 34 comprise four tilting EPUs 54 mounted fore of the wings and four fixed EPUs 56 mounted aft of the wings. As shown in Figure 4, each EPU is labelled from EPU1 to EPU8. EPU1 to EPU4 are tilting EPUs 54, and EPU5 to EPU8 are fixed EPUs 56.

[0162] It will be appreciated that each of the tilting EPUs 54, EPU1 to EPU4, is fixed to a respective fore pylon portion 90 (as shown in Figure 2). Each tilting mechanism 35 is operable to tilt the rotor 33 connected to the respective tilting EPU 54 relative to the pylon 40 (and thereby the airframe 24) between a vertical flight mode in which the rotors 33 deliver thrust which is parallel to the yaw axis 18 and a convention flight mode in which the rotors 33 deliver thrust which is parallel to the roll axis 16 of the eVTOL aircraft 10.

[0163] Each of the fixed EPUs 56, EPU5 to EPU8, is fixed to a respective aft pylon portion 92. The fixed EPUs 56 are fixed relative to the pylons 40 (and thereby the airframe 24) such that rotors

[0164] 33 connected to the fixed EPUs 56 deliver thrust which is parallel to the roll axis 16 of the eVTOL aircraft 10 only. The arrangement of the EPUs 34 relative the planes 20,22 of the eVTOL aircraft 10 is also illustrated in Figure 4. EPU1 and EPU2 are fixed to the airframe 24 on the left side 104 of the left / right plane 20 and the fore side 100 of the fore / aft plane 22. EPU3 and EPU4 are fixed to the airframe 24 on the right side 106 of the left / right plane 20 and the fore side 100 of the fore / aft plane 22. EPU5 and EPU6 are fixed to the airframe 24 on the left side 104 of the left / right plane 20 and the aft side 102 of the fore / aft plane 22. EPU7 and EPU8 are fixed to the airframe 24 on the right side 106 of the left / right plane 20 and the aft side 102 of the fore / aft plane 22.

[0165] Figure 4 also shows schematically the EPUs 34 to which each of the busbars 32 is connected. It will be appreciated that the location of the busbars 32 shown in Figure 4 are merely schematic and may not represent the physical location of the busbars 32 within the airframe 24. The eight busbars 32 are labelled busbar one, BB1, to busbar eight, BB8. Busbar BB3 is connected to EPU5 and EPU3, and busbar BB7 is connected to EPU2 and EPU8. Thereby, busbar BB3 and busbar BB7 are connected to EPUs 34 on different sides of the left / right plane 20. Busbars BB1, BB2, BB4, BB5, BB6, and BB8 (i.e. a majority of the busbars 32) are connected to EPUs

[0166] 34 which are on the same side of the left / right plane 20.

[0167] Although a specific architecture, in terms of wing configuration, and the number and arrangement of EPUs and batteries, of eVTOL aircraft is illustrated herein, it will be clear that the techniques for equalising power demand on batteries under different scenarios, and the regimes of operation, described herein apply equally to other architectures that comprise different configurations of wings, and numbers and arrangements of EPUs and batteries.

[0168] The four titling EPUs 54 are identical, so only one titling EPU 54 will be described for the sake of brevity. The tilting EPU 54 (shown in Figure 5) comprises: an electric motor 60, a dual lane inverter apparatus 64, and an EPU failure sensor 66. The electric motor 60 is a three phase AC electric motor 60 and comprises a pair of three-phase windings 68 and a motor rotor 70. The windings 68 are electrically isolated from one another and can cause rotation of the motor rotor 70 together or independently of one another. The maximum power rating of the electric motor 60 in this example is 452kW.

[0169] The dual lane inverter apparatus 64 comprises: a pair of busbar connections 76, a pair of winding connections 78, a power command connection 80, a pair of inverters 82, and a voltage controller 83. Each of the busbar connections 76, winding connections 78, and power command connection 80 comprises an electric cable. Each busbar connection 76 comprises a protection device 84. The protection device 84 is provided to protect the electric components of the eVTOL aircraft 10 in the event of a fault.

[0170] Each inverter 82 comprises a voltage sensor 86, and an inverter controller 89. The voltage sensor 86 measures a voltage at an input of the inverter 82. The inverter controller 89 controls the inverter 82 in a known way in response to commands received from the voltage controller 83 (discussed below in more detail). Each inverter 82 in this example is a bidirectional inverter, which supports both drive and regenerative modes of operation.

[0171] Each inverter 82 has a maximum power rating of 360kW. Therefore, the maximum power rating of each inverter 82 is approximately 80% of the maximum power rating of the electric motor 60. Inverter 82 maximum power ratings of no less than 60% and no greater than 85% of the maximum power rating of the electric motor 60 are also envisaged.

[0172] The voltage controller 83 controls the pair of inverters 82 by sending commands to the inverter controllers 89, as will be discussed below in more detail. The EPU failure sensor 66 detects the failure of the tilting EPU 54.

[0173] The four fixed EPUs 56 are similar to the tilting EPUs 54, aside from not being connected to a tilting mechanism 35 (discussed in more detail below).

[0174] The eight rotors 33 comprise five bladed rotors 63 and four bladed rotors 65 (shown in Figure 2). The five bladed rotors 63 each comprise five blades 72 and a hub 74. The blades 72 are connected to the hub 74, and the hub 74 is connected to the electric motor rotor 70 of one of the tilting EPUs 54 such that rotation of the electric motor rotor 70 causes rotation of the hub 74 and the blades 72.

[0175] The four bladed rotors 65 each comprise four blades 72, and a hub 74. The blades 72 are connected to the hub 74, and the hub is connected to the electric motor rotor 70 of one of the fixed EPUs 56. Additionally, the four bladed rotors 65 comprise a stowage mechanism (not shown) for moving the blades 72 between a stowed position and a deployed position (shown in Figure 1).

[0176] Each tilting mechanism 35 is connected to one of the tilting EPUs 54. The function of tilting mechanisms 35 is described below.

[0177] The electric connections between the electric components of the eVTOL aircraft 10 will now be discussed with reference to Figure 6. The eight batteries 30 are labelled battery one, Bl, to battery eight, B8. The electric motor 60 of each EPU 34 is labelled in accordance with the respective EPU number. For example, the electric motor 60 of EPU4 is labelled M4. The inverters 82 of each EPU 34 are labelled in accordance with the respective EPU number, and each inverter 82 is labelled either ‘a’ or ‘b’. For example, the pair of inverters 82 comprised by EPU4 are labelled I4a and I4b. The arrangement illustrated in Figure 6 reflects the positions of the EPUs shown in Figure 4.

[0178] The flight controls 26 are connected to the flight control computer 28. The flight control computer 28 is connected to: the voltage controller 83 of each EPU 34 via the power command connection 80 of each EPU 34, the EPU failure sensor 66 of each EPU 34, the stowage mechanism of each of the four-bladed rotors 65, and the tilting mechanism 35.

[0179] As can be appreciated from Figure 6, the EPUs are connected in a loop configuration, via the respective arrangements of inverters 82, motors 60 and busbars 32. Any failure of an inverter 82, a motor 60 or a busbar 32, effectively breaks the loop configuration. A failure of a battery 30 as such does not necessarily break the loop configuration, unless the battery 30 failure also causes the respective busbar 32 to fail, breaking the connection between the two respective inverters 82. In any case, the loop configuration provides resilience in the case of such failures, as will become apparent from the following description. Each battery 30 is connected to a busbar 32, and so feeds two EPUs 34; or, more particularly in this example, each battery feeds one lane and respective set of windings of each of two EPUs 34. For example, battery B4 is connected to busbar BB4. In this example, circuitry (not shown) comprising an arrangement of one or more diodes is provided between each battery 30 and its respective busbar 32. The diodes are arranged to inhibit energy flowing from the busbars 32 back to the battery 30. This may improve the fault tolerance of the batteries 30 by reducing the likelihood of high currents, which could damage the batteries 30, flowing back into the batteries 30 during a fault event. Additionally, the circuitry prevents regenerated energy (discussed below in more detail) from reaching the battery 30. In other examples, the circuitry may be omitted.

[0180] Each busbar 32 in this example is connected to a fixed EPU 56 and a tilting EPU 54. For example, busbar BB4 is connected to EPU3 and EPU7. Each busbar 32 is connected to only two EPUs 34, in this example.

[0181] Each inverter 82 is connected via a respective busbar connection 76 to a respective busbar 32 and thereby connected to a respective battery 30 via the respective busbar 32. For example, inverter I4a is connected to battery B5 via busbar BB5. Each inverter 82 is also connected via one of the pair of winding connections 78 to a respective winding 68.

[0182] Each voltage controller 83 is connected to the inverter controllers 89 of the pair of inverters 82, the voltage sensors 86, and the power command connection 80 of the EPU 34 comprising the respective voltage controller 83. Thereby, the voltage controller 83 is arranged to control the level of power that is delivered from the respective battery 30 connected to the inverter 82, to the winding 68 connected to the inverter 82, by sending commands to the inverter controller 89. For example, the voltage controller 83 of EPU4 is arranged to send commands to the inverter controller 89 of inverter I4a and thereby control the level of power delivered from battery B5 via busbar BB5 to the winding 68 of motor M4.

[0183] The operation of the eVTOL aircraft 10 will now be described. To aid comprehension, the different regimes which the voltage controllers 83 employ will be explained in the context of different scenarios initially without reference to decision logic applied by the voltage controllers 83 to determined which regime to employ (Figures 7 to 15). The decision logic 201, which each voltage controller 83 follows to determine which regime to operate in, will be described with reference to Figure 16.

[0184] Firstly, a regime will be described in which, in response to there being a difference in the voltages of the batteries 30, the level of power of the inverter 82 with the higher voltage is increased and the level of power of the inverter 82 with the lower voltage is reduced to balance the voltages. This regime can adjust the levels of power in a manner designed to balance the voltages in a reasonable time. For instance, in the example described with reference to the flow chart in Figure 7, the ratio of the levels of power is varied proportionally to the difference in the voltages, and the regime is referred to as a proportional regime. However, in other examples, this regime may employ other approaches for determining the magnitude of the change in the levels of power whilst still increasing the level of power of the inverter 82 with the higher voltage and reducing the level of power of the inverter 82 with the lower voltage to balance the voltages. For example, if there is a voltage difference, a predefined ratio of 0.8: 1.2 could be used until voltage balancing is achieved. The proportional regime will be described in the context of a flight in which a component failure does not occur, with reference to the flow chart in Figure 7. In this example, the voltage controllers 83 only operate in the proportional regime.

[0185] During the flight operation (e.g., take-off), the pilot inputs a flight command 101 which is communicated to the flight control computer 28. The flight control computer 28 comprises logic which converts the flight command into eight power commands 103. Each power command indicates a demanded level of power that is to be delivered to a respective motor 60. For example, each power command may be a demanded speed of the respective rotor 33. The power commands are communicated 105 to the voltage controllers 83 of the EPUs 34. For the take-off, the power commands indicate a demanded level of power greater than the pre-take off level of power of the electric motors 60, so are commands to increase the power of the electric motors 60. For the take-off, the tilting EPUs 54 are in the vertical flight mode.

[0186] The operation of only a single EPU 34 will now be described. However, it will be appreciated that the other EPUs 34 operate in a similar manner across the loop configuration to equalise the power demand (and respective discharge rate) of each battery 30. The voltage controller 83 of the EPU 34 receives 105 its respective power command. Additionally, the voltage controller 83 receives 105 a pair of measurements from the voltage sensors 86 of the input voltage to each of the pair of inverters 82. The voltage controller 83 determines 107 a ratio of the levels of power of each inverter 82. This is performed by varying a previous ratio of the levels of power proportionally to the difference in the voltages indicated by the pair of measurements. Figure 8 shows an example, immediately after commencing takeoff, in which the voltage measurements are equal, so the determined ratio remains at an initialised ratio of 1 :1. An example of varying the ratio when there is a voltage difference is described below with reference to Figures 9a to 9c.

[0187] Once the ratio has been determined, the voltage controller 83 apportions 109 the level of power demand, indicated by the power command, between the inverters 82. The level of power demand is apportioned in accordance with the ratio to determine a level of power for each inverter 82. For example, in Figure 8, the level of power demand for EPU3 (252kW) is apportioned in accordance with the determined ratio of 1 : 1 such that the level of power for inverter I3a is 134kW and the level of power for inverter I3b is 134kW.

[0188] The voltage controller 83 then sends commands to the respective inverter controllers 89 indicating the level of power of each inverter 82. In response, each inverter controller 89 then controls 111 their respective inverter 82 such that the determined level of power for each inverter 82 is achieved and the demanded level of power is delivered to the windings 68.

[0189] The same process is performed by each EPU 34, such that the steady state situation shown in Figure 8 is achieved. As illustrated in Figure 8, each of the fore motors (M1-M4) has a power demand of 252kW and each of the aft motors (M5-M8) has a power demand of 148kW. The inverters 82 for each EPU 34 share the demand 1 : 1, such that each of the fore motors (M1-M4) draws 126kW from each respective inverter 82, and each of the aft motors (M5-M8) draws 74kW from each respective inverter 82. Given the loop configuration of the system, each battery 30 delivers 200kW and is in balance, in terms of power demand and discharge rate, with each other battery 30.

[0190] In this scenario, there is no difference between the voltage measurements and therefore the ratios remain at their initialised values of 1 :1. This results in the power level of each inverter 82 within an EPU 34 being equal. The ratios of 1 : 1 may last only for the starting instance of the take-off when all the batteries 30 are fully charged, and thereafter the ratios are dynamically altered during flight, as will now be described.

[0191] Figure 9a shows a scenario during a take-off where the voltages are different. In Figure 9a, the flight control computer 28 has increased the level of power demand for EPU7 from 148kW to 200kW (e.g. to counteract wind forces and balance the eVTOL aircraft 10). This results in the level of power of battery B4 increasing from 200kW to 226kW. The voltage of each battery 30 will be related to the level of power of the battery 30. Batteries 30 with a higher level of power demand (and thereby discharge rate) will have a lower voltage than batteries 30 with a lower level of power (and thereby a lower discharge rate) due to voltage droop effects. If left unchecked, batteries 30 would continue to discharge at uneven rates, with associated premature loss of utility of the eVTOL aircraft 10, when one (or too many) batteries 30 become depleted. As the level of power demand from battery B4 is greater than from battery B3, the voltage of battery B4 will be lower than the voltage of B3. As a result, the voltage measurements received by the voltage controller 83 of EPU3 will show a voltage difference between the input voltages of inverters I3a and I3b, with the input voltage to I3b being lower than the input voltage to I3a.

[0192] The voltage controller 83 determines a new ratio of the levels of power of each inverter 82, by varying the previous ratio (1 :1) of the levels of power proportionally to the difference in the voltages indicated by the pair of measurements. The voltage controller 83 varies the ratio to increase the proportion of the EPU’s level of power demand met by inverter I3a, which has the higher voltage, and reduce the proportion of the EPU’s level of power demand met by inverter I3b, which has the lower voltage. This is shown in Figure 9b, in which the level of power of I3a is 139kW and the level of power of inverter I3b is 113kW. This results in the difference in the levels of power (and thereby discharge rates) of batteries B3 and B4 reducing and becoming more balanced, as shown in Figure 9b in which batteries B3 and B4 each have a level of power of 213kW.

[0193] As shown in Figure 9b, the levels of power (and thereby the voltages due to the voltage droop effect) of battery B2 is not equal with the level of power of B3 (200kW v 213kW). Additionally, the levels of power (and thereby the voltages due to the voltage droop effect) of battery B5 is not equal with the level of power of B4 (200kW v 213kW). As a result, a chain reaction of changing ratios is commenced, the start of which is described below. Due to the levels of power difference between battery B2 and battery B3, the voltage measurements received by the voltage controller 83 of EPU5 will show a voltage difference between the input voltages of inverters I5a and I5b, with the input voltage to I5b being lower than the input voltage to I5a. In response, the voltage controller 83 of EPU5 determines a new ratio of the levels of power of each inverter 82, by varying the previous ratio of the levels of power (1 : 1) proportionally to the difference in the voltages indicated by the pair of measurements. The voltage controller 83 of EPU5 varies the ratio to increase the proportion of the EPU’s level of power demand met by inverter I5a, which has the higher voltage, and reduce the proportion of the EPU’s level of power demand met by inverter I5b, which has the lower voltage. This results in the scenario shown in Figure 9c. Inverter I5a has a level of power of 80.5kW and inverter I5b has a level of power of 67.5kW. Accordingly, the level of power of batteries B2 and B3 will be 206.5kW and in balance.

[0194] Due to the levels of power difference between battery B4 and Battery B5, the voltage measurements received by the voltage controller 83 of EPU7 will show a voltage difference between the input voltages of inverters I7a and I7b, with the input voltage to I7a being lower than the input voltage to I7b. In response, the voltage controller 83 of EPU7 determines a new ratio of the levels of power of each inverter 82, by varying the previous ratio (1 : 1) of the levels of power proportionally to the difference in the voltages indicated by the pair of measurements. The voltage controller 83 of EPU7 varies the ratio to increase the proportion of the EPU’s level of power demand met by inverter I7b, which has the higher voltage, and reduce the proportion of the EPU’s level of power demand met by inverter I7a, which has the lower voltage. This results in the scenario shown in Figure 9c. Inverter I7b has a level of power of 106.5kW and inverter I7a has a level of power of 93.5kW. Accordingly, the level of power of batteries B4 and B5 will be 206.5kW and in balance. The levels of power of batteries Bl and B6 which are adjacent to batteries B2 and B5 will be 200kW. Therefore, there will be a voltage difference in the measurements received by EPU1 and EPU4. EPU1 and EPU4 will react in a similar manner to that described above for EPU5 and EPU7 to balance the voltages. This chain reaction continues to propagate through the remaining EPUs (EPU2, EPU6, and EPU8), which balances the level of power provided by each of the batteries 30.

[0195] Referring now to Figure 10, once the take-off has been completed, the flight control computer 28 controls the tilting mechanisms 35 to tilt 121 the rotors 33 of the tilting EPUs 54 from the vertical flight mode to the conventional flight mode. The flight control computer 28 sends 123 new power commands to the voltage controllers 83 of the EPUs 54, which indicate zero power for the fixed EPUs 56. In response to the new power commands, the inverters 82 are controlled to deliver zero power to the electric motors 60 (M5-M8) of the fixed EPUs 56. The flight control computer 28 then controls the stowage mechanism of each of the four bladed rotors 65 to move 125 the blades 72 from their deployed position to their stowed position. As each battery 30 serves a fore and an aft motor, each battery 30 continues to be used to drive a fore motor even after the aft motors have been powered down.

[0196] Prior to performing the landing, the flight control computer 28 controls the stowage mechanism of each of the four bladed rotors 65 to move 127 the blades 72 from their stowed position to their deployed position. To perform the landing, the flight control computer 28 controls the tilting mechanisms 35 to tilt 129 the rotors 33 of the tilting EPUs 54 from the conventional flight mode to the vertical flight mode and the flight control computer 28 sends 131 further power commands to the voltage controllers 83 of all the EPUs 34 to increase the power of all the EPUs 34. During the transitions between vertical flight mode and conventional flight mode, additional flight commands from the flight controls 26 and power commands from the flight control computer 28 may be sent to maintain the stability of the eVTOL aircraft 10.

[0197] A failure regime will now be described in the context of the eVTOL aircraft’s 10 response to the failure of busbar BB4 or battery B4 with reference to Figures 11 and 12. In the failure regime, one of the inverters 82 is no longer able to draw power due to the failed component. In response, the voltage controller 83 determines a ratio of 1 :0 such that the demanded power of the EPU 34 is provided by the inverter 82 which is still able to draw power. Figure 11 illustrates power demand on the batteries immediately after a component failure while Figure 12 illustrates the power demand once balancing has taken effect. Employing the failure regime may enable an EPU 34 to maintain its level of power in the event of a component failure which prevents power being delivered through one of its inverters 82. However, as will be described below, using the failure regime may place increased power demands on the batteries 30 adjacent to the failed component. The loop configuration provides a means for redistributing this increased power demand and balance the batteries 30. In this example, the voltage controllers 83 of the EPUs 34 connected to the failed busbar BB4 or failed battery B4 (EPU 3 and EPU7) operate in the failure regime, whilst the voltage controllers 83 of the remaining EPUs (EPUs 1, 2, 4, 5, 6, and 8) operate in the proportional regime described previously with reference to Figures 9a to 9c. After the failure of busbar BB4 or battery B4, power is no longer delivered from battery B4 to inverters I3b and I7a. This results in the input voltages measured by the voltage sensors 86 of the affected inverters, I3b and I7a, connected to busbar BB4 dropping to 0V. Batteries B3 and B5 and busbars BB3 and BB5 have not failed, so power is still delivered to inverters I3a and I7b. As power is still being delivered, the input voltages measured by the voltage sensors 86 comprised by the non-affected inverters, 13 a and I7b, will be greater than 0V. In response to receiving a 0V measurement (described below in more detail), the voltage controllers 83 of EPU3 and EPU7 operate in the failure regime. In the failure regime, the inverters 82 determine a ratio of 1 :0 such that the demanded power of EPU3 and EPU7 is provided by the inverter 82, which is still able to draw power. Specifically, the non-affected inverters I3a and I7b, with the non-zero voltage measurements, provide the demanded power to EPU3 and EPU7. As shown in Figure 11, this results in the level of power of I3a being 220kW to meet EPU3’s level of power demand and the level of power of inverter I7b being 180kW to meet EPU7’s level of power demand.

[0198] The voltage controllers 83 of the EPUs (EPUs 1, 2, 4, 5, 6 and 8) not connected to busbar BB4 operate in the proportional regime, as described with reference to Figures 9a to 9c, to reduce the difference in the voltages indicated by the measurements and balance the level of power provided by each of the batteries 30. The increased level of power of the non-affected inverters, I3a and I7b, results in an increase in the level of power (and thereby discharge rate) from batteries B3 and B5 to which the non-affected inverters I3a and I7b are connected. The level of power delivered by battery B3 increases to 310kW, and the level of power delivered by battery B5 increases to 290kW. The levels of power of the batteries B2 and B6 are lower than those for batteries B3 and B5 (200kW compared with 310kW or 290kW), which (due to the voltage droop effect) results in a difference in the input voltages for the inverters 82 of EPU5 and EPU4. These differences in input voltages result in the ratios for EPU5 and EPU4 changing, and subsequently result in a chain reaction of changing ratios for EPUs 1, 2, 6 and 8. This chain reaction balances the level of power provided by each of the batteries 30 (228.6kW), as shown in the steady state situation of Figure 12.

[0199] A regenerative regime will now be described in the context of an inverter 82 failure, shown in Figure 13. Figure 13 illustrates the power demand once balancing has taken effect. The regenerative regime may enable the batteries 30 to be balanced to a greater extent in scenarios in which the proportional regime alone may not be able to fully balance the batteries 30 (e.g. as illustrated in Figure 9b). This may be achieved by effectively redistributing power from an underutilised battery 30, through the EPU 34 to which the battery 30 is directly connected, to an EPU 34 to which the battery 30 is not directly connected, as will be described below. When operating in the regenerative regime, the voltage controllers 83 determine a negative ratio such that the voltage controllers 83 control individual inverters 32 to operate in a regenerative mode to convert AC power harvested from the winding 68 to which it is connected within an EPU 34, where the electric motor rotor 70 of the EPU 34 is driven by the other winding 68 of the EPU 34. As has already been stated, regenerative operation requires the invertors to be bidirectional.

[0200] In the example of Figure 13, the voltage controller of EPU3 operates in the failure regime described above with reference to Figures 11 and 12, and the voltage controllers 83 of EPU5, EPU6, EPU7, and EPU8 operate in the regenerative regime. The voltage controllers 83 of EPU 1, EPU2, and EPU4 operate in the proportional regime described with reference to Figures 9a to 9c. It should be noted that the voltage controllers 83 of EPU1, EPU2, and EPU4 operate at the extremes of the proportional regime by employing ratios of 0: 1 such that zero power is delivered through inverters 82 which are able to have power delivered through them. The inventors have identified that, whilst potentially counter-intuitive, doing this may enable more accurate battery 30 balancing in certain circumstances, such as the inverter 82 failure scenario shown in Figure 13.

[0201] After the failure of inverter I3a of EPU3, the voltage measurement for inverter I3a drops to 0V. In response, the voltage controller 83 of EPU3 operates in the failure regime and determines a ratio of 1 :0 such that the level of power of the non-failed inverter I3b is equal to the level of power demand of EPU3 (252kW). The non-failed inverter I3b is then controlled to deliver the demanded level of power to the winding 68 connected to the non-failed inverter I3b.

[0202] The increased level of power of the non-failed inverter I3b results in an increase in the discharge rate from the battery B4 to which the non-failed inverter I3b is connected. This results in a reduction in the input voltage to the other inverter I7a to which battery B4 is connected via busbar BB4. The voltage controller 83 of EPU7 receives the pair of measurements from the voltage sensors 86 of inverters I7a and I7b. The voltage controller 83 of EPU7 then operates in the regenerative regime and determines a ratio of the levels of power of each inverter 82, which reduces the difference in the voltages indicated by the measurements. In contrast to the previous example, this ratio is negative such that when the voltage controller 83 apportions the level of power demand indicated by the power command between the inverters 82 in accordance with the ratio, the level of power for I7a is negative (-52kW).

[0203] In response, the inverters I7a and I7b are controlled such that the winding 68 to which inverter I7a is connected regeneratively harvests energy from the electric motor M7 and supplies this regeneratively harvested energy to busbar BB4. Inverter I7b is controlled to deliver the demanded level of power (148kW) plus the regeneratively harvested power (52kW) and thereby has a level of power of 200kW.

[0204] Similarly, the decreased level of power of inverter I3a results in a reduction in the discharge rate from battery B3. The voltage controller 83 of EPU5 operates in the regenerative regime and in a corresponding manner to the voltage controller 83 of EPU7 described above resulting in the level of power of I5a being negative (-52kW) and regenerating power from motor M5 to busbar BB2.

[0205] The behaviour of the voltage controllers 83 of EPU5 and EPU7 subsequently results in a chain reaction of changing ratios for the other EPUs (EPUs 1, 2, 4, 6 and 8) which balances the level of power (200kW) provided by each of the batteries 30, as shown in the steady state scenario of Figure 13. As noted above, EPU1, EPU4 and EPU2 operate at the extremes of the proportional regime, employing ratios of 0: 1. Using EPU1 as an example, inverter Ila draws OkW from battery Bl despite there being no faults in battery Bl, busbar BB1, or inverter Ila which would prevent power being delivered. Meanwhile, inverter I lb provides the full 252kW demanded by the flight control computer 28 for EPU1. While this mode of operation may apper counterintuitive, it leads to balancing the level of power across all batteries.

[0206] A scenario employing the proportional regime and the regenerative regime will now be described in the context of an EPU failure (Figures 14 and 15). Figure 14 illustrates power demand on the batteries immediately after the EPU failure, while Figure 15 illustrates the power demand once balancing has taken effect. Unlike in the previous failure scenarios, the failure of an EPU 34 means the original power levels of one or more of the EPUs 34 must be adjusted if the overall power demand is to be maintained. As a result, the flight control computer 28 sends power commands to various EPUs 34 to compensate for the loss of thrust from the failed EPU1 and balance the eVTOL aircraft 10. This results in different demands being placed on each of the batteries 30. These different demands are ultimately balanced by the control regimes as will be described below.

[0207] In this example, the voltage controller 83 of EPU7 operates in the regenerative regime described above with reference to Figure 13, and EPUs 2 to 6, and 8 operate in the proportional regime described above with reference to Figures 9a to 9c.

[0208] The failure of EPU1 results in a reduced demand being placed on batteries Bl and B2. As shown in Figure 14, the power into invertors Ila and lib reduces to zero. As a result, the power demand and respective discharge rates of batteries Bl and B2 reduce compared to the other the batteries 30. The reduced power demand on batteries Bl and B2, as a result of voltage droop effects, leads to a relatively elevated voltage on respective busbars BB 1 and BB2 compared to when EPU1 was operational and drawing power. The following description explains how the system reacts to these changes.

[0209] The EPU failure sensor 66 associated with EPU 1 detects the failure of EPU 1 and sends a message to the flight control computer 28. In response, the flight control computer 28 sends updated power commands to the remaining EPUs (EPU2-EPU8), as shown in Figure 14. The flight control computer 28 sends an updated power command to decrease the power of EPU7 (-5.2kW, which represents a deceleration of EPU7) and EPU8 (44.58kW), which are located on the opposite side of the fore / aft plane 22 and left / right plane 20 to EPU1, The flight control computer 28 sends updated power commands to increase the power of EPU2 (416.54kW), which is on the same side of the fore / aft plane 22 and the left / right plane 20 as EPU1, EPU3 (416.54kW) and EPU4 (309.2kW), which are both on the same side of fore / aft plane 22 as EPU1, EPU 5 (321.26kW) and EPU6 (220.24kW), which are both on the same side of the left / right plane 20 as EPU1. The updated power commands are sent to compensate for the loss of thrust from EPU1 and to balance the moments of the thrust generated by the rotors 33 of the remaining EPUs (EPU2-EPU8).

[0210] To aid in comprehension, in this example the measurements of the voltages received immediately before receiving the updated power commands are equal, and the voltage controllers 83 initially maintain the ratios at 1 :1. However, in practice, it will be appreciated that, during a flight mission, the voltage ratios are constantly changing as conditions, and respective power demands on each EPU 34, vary. As such, in most practical scenarios, the detected voltage ratio for any particular EPU 34 may not be exactly 1 :1 prior to an EPU 34 failure. However, in these cases, the eVTOL aircraft 10 will operate in a corresponding manner to that described above with reference to starting ratios of 1 : 1.

[0211] Each voltage controller 83 of the remaining EPUs (EPU2-EPU8) initially apportions the updated level of power demand indicated by their updated power command between their pair of inverters 82 in accordance with the maintained ratio of 1 : 1 to determine a new level of power for each of their inverters 82. Each inverter 82 is then controlled to deliver the new level of power. This results in the level of power drawn from each battery 30 being different, as shown in Figure 14: Bl (110.12kW), B2 (160.64kW), B3 (368.92kW), B4 (205.68kW), B5 (152kW), B6 (176.9kW), B7 (230.58kW) and B8 (318.4kW). Due to the voltage droop effect, this leads to differences in the voltages of the batteries 30.

[0212] In response to receiving updated pairs of voltage measurements, which indicate the differences in voltages described above, the voltage controllers 83 of the remaining EPUs (EPU2-EPU8) determine updated ratios of the levels of power of each inverter 82, which are to reduce the differences in the voltages indicated by the updated pairs of measurements.

[0213] Once the updated ratios have been determined, the voltage controllers 83 apportion the updated level of power demands indicated by the updated power commands between the inverters 82, in accordance with the updated ratios, to determine an updated level of power for each inverter 82. Each inverter 82 is then controlled to deliver the updated level of power. The voltage controller 83 continues the above steps, in one or more iterations, until the steady state situation shown in Figure 15 is achieved. As shown in Figure 15, this results in the balancing of the level of power provided by each of the batteries 30. More particularly, when EPU1 fails, and EPU7 and EPU8 are powered down or off, the five other EPUs 34 effectively compensate for loss of thrust from EPU1, EPU7 and EPU8. The consequences of this are significant, where the initial demands placed on plural EPUs (EPU2, EPU3) exceeding 400kW lead to demands on two of the batteries (B3 and B8) exceeding 300kW. After balancing across the system, the demands on the batteries 30 are reduced to approximately 215kW.

[0214] In order to determine which regime to operate in, each voltage controller 83 employs decision logic 201, which will now be described with reference to Figure 16. The decision logic 201 commences with the voltage controller 83 determining 203 if the difference in voltages is below a noise threshold. In this example, the noise threshold is IV. The noise threshold may reduce the likelihood of the voltage controller 83 changing the ratio when the voltages are already effectively balanced. If the voltage difference is less than the noise threshold, then the voltage controller 83 does not alter the ratio and returns to the start of the decision loop 201, effectivly waiting until a more significant voltage difference occurs.

[0215] If the voltage difference is greater than the noise threshold, the voltage controller 83 determines 205 if the flight control computer (FCC) 28 has demanded zero power for an EPU 34. If this is the case, then the voltage difference is intentional, and the voltage controller 83 returns to the start of the decision loop 201. If this is not the case, then the voltage controller 83 proceeds to determine 207 if either of the voltages equals zero.

[0216] If either voltage equals zero, then a failure has occurred, and the voltage controller 83 proceeds to operating 209 in the failure regime described above with reference to Figures 11 and 12 in order to maintain the level of power of the EPU 34. After operating 209 in the failure regime, the voltage controller 83 returns to the beginning of the decision loop 201.

[0217] If both of the voltages are greater than zero, the voltage controller 83 operates in the proportional regime 211 described above with reference to Figures 9a to 9c. The voltage controller 83 then determines 213 if the voltage difference is below the noise threshold. If the voltage difference is below the noise threshold, then the voltage controller 83 returns to the start of the decision loop 201.

[0218] If the voltage difference is above the noise threshold (which means the voltages and thereby the batteries’ 30 discharge rates are still not sufficiently balanced), then the voltage controller 83 performs a series of determinations to decide if the regenerative regime described above with reference to Figure 13 should be employed. The voltage controller 83 determines 215 if the ratio is equal to 0: 1, which means one of the inverter’s 82 level of power is zero whilst the other inverter 82 is providing the full power demanded by the flight control computer 28 for the EPU 34.

[0219] The ratio not equalling 0:1 means that further alteration of the ratio within the proportional regime is possible to further balance the voltages. Conversely, the ratio equally 0:1 means that the proportional regime has reached its limit, and no further alteration of the ratio is possible in the proportional regime to further balance the battery 30. In this case, switching to the regenerative regime may enable better voltage balancing.

[0220] If the ratio does not equal 0:1, the voltage controller 83 returns to the start of the decision loop 201. Thereby, the voltage controller 83 is able to iterate around the decision loop to employ multiple iterations of the proportional regime.

[0221] If the ratio equals 0: 1, the voltage controller 83 determines 217 if the maximum power rating of the inverter 82 with the higher input voltage (which is the inverter 82 which would be required to draw both the demanded level of power of the EPU 34 and the power to be regenerated) is greater than the level of power demand for the EPU 34. This is to ensure that the inverter 82 has sufficient power capacity to provide the power of the EPU 34 and additional power for regeneration. If the maximum power rating is insufficient, then the voltage controller 83 returns to the start of the decision loop 201.

[0222] If the maximum power rating is sufficient, then the voltage controller 83 determines 219 if the higher input voltage is greater than a minimum voltage. The minumum voltage is set to a level which corresponds to a battery 30 discharge rate above which the likelihood of thermal runaway events becomes unnacceptable. As a result, this determination is to prevent the voltage controller 83 drawing additional power through the inverter 82 with the higher input voltage in scenarios where this may lead to a high discharge rate and associated higher likelihood of a thermal runaway in the respective battery 30. If the voltage is less than the minimum voltage, then the voltage controller 83 returns to the start of the decision loop 201.

[0223] If the voltage is greater than the minimum voltage, the voltage controller 83 determines 221 if the voltage difference is below the noise threshold. Although potentially superfluous on the first pass through the decision loop 201, this determination may enable multiple iterations of the regenerative regime to be employed as described below. If the voltage difference is below the noise threshold, then the voltage controller 83 returns to the start of the decision loop 201.

[0224] If the voltage difference is above the threshold, then the voltage controller 83 operates 223 in the regenerative regime described with above reference to Figure 13. After operating 223 in the regenerative regime, the voltage controller 83 returns to the previous determination 217 of if the maximum power rating of the inverter 82 is sufficient. The voltage controller 83 then continues through the previously described steps of determining 219 if the input voltage is greater than the minumum voltages, determining 221 if the voltage difference is below the noise threshold, and operating 223 in the regenerative regime. In this way, whilst a significant voltage difference remains, and the inverter 82 with the higher voltage is still in an acceptable state to operate in the regenerative regime, the voltage controller 83 will operate in the regenerative regime for multiple iterations.

[0225] In the above example described with reference to Figures 13 to 15, the voltage controllers 83 are able to operate in the regenerative regime. Accordingly, the inverters 82 are bidirectional inverters. However, in other examples, the inverters 82 are unidirectional inverters and the voltage controller 83 operates decision logic 301 which omits the steps related to the regenerative regime. Such a decision logic 301 is shown in Figure 17. The decision logic 301 of Figure 17 is identical to the decision logic 201 of Figure 16, except that once the voltage controller 83 operates in the proportion regime, the voltage controller 83 returns to the start of the decision logic 301. The steps 203, 205, 207, 209, and 211 of Figure 17 are the same as the steps 203, 205, 207, 209, and 211 of Figure 17.

[0226] In the above examples described with reference to Figures 1 to 17, the pair of measurements are of the input voltage to each of the inverters 82. These input voltages are indicative of the discharge rate of the battery 30 to which each of the inverters 82 in connected. In other examples, each measurement may be of the current flowing between each battery 30 and the respective busbar 32 to which each battery 30 is connected. The current would also be indicative of the discharge rate of the respective battery 30. Therefore, in a more general sense, it may be said that the pair of measurements are indicative of the discharge rate of the respective battery 30 connected to each inverter 82.

[0227] In the above examples described with reference to Figures 1 to 17, inverters 82 are used to control the level of power delivered to the windings 68 of the electric motors 60. In other examples, the electric motors may be DC and other forms of power control units may be used, such as a power transistor. Therefore, in a more general sense, it may be said that an appropriate power control unit is used to control the level of power delivered to the windings 68 of the electric motor 60. In the above example described with reference to Figure 13, after an inverter 82 failure, the voltage controller 83 determines a ratio of 1 :0 such that the level of power of the non-failed inverter lib is equal to the level of power demand indicated by the power command. However, in other examples, the voltage controller 83 may control the non-failed inverter lib such that the level of power of the non-failed inverter 82 is equal to the maximum power rating of the non-failed inverter lib. In this example, the flight control computer 28 may send updated power commands to the other EPUs EPU2-EPU8 which do not comprise the failed inverter Ila to balance the moment of the thrust generated by EPU1, which comprising the failed inverter Ila.

[0228] In the above example described with reference to Figures 11 and 12, after a busbar 32 or battery 30 failure, the voltage controllers 83 determine a ratio of 1 :0 such that the level of power of the non-affected inverters I3a I7b is equal to the level of power demand indicated by the power command. However, in other examples, the voltage controllers 83 may control the non-affected inverter I3a I7b such that the levels of power of the non-affected inverters I3a I7b is equal to the maximum power ratings of the non-affected inverters I3a I7b. In this example, the flight control computer 28 may send updated power commands to the EPUs (EPU1,2,4,5,6 and 8) not connected to busbar four BB4 to balance the moment of the thrust generated by the EPUs (EPU3 and EPU7) connected to the failed busbar BB4 or the failed battery B4.

[0229] An alternative connection arrangement for the busbars 32, EPUs 34 and batteries 30 is shown in Figure 18.

[0230] Although in the above examples described with reference to Figures 1 to 17, the eVTOL aircraft 10 comprises eight batteries 30 and eight EPUs 32. It will be appreciated that in other examples, the eVTOL aircraft 10 may comprise a different number of these components. For example, the eVTOL aircraft may comprise six batteries and six EPUs.

[0231] In the above examples described with reference to Figures 1 to 17, the eVTOL aircraft 10 comprises eight EPUs 34 (four fixed EPUs 56 and four tilting EPUs 54), and the EPUs 34 are located in a particular arrangement with respect to the left / right plane 22 and the fore / aft plane 22. In other examples, the eVTOL aircraft 10 may comprise a different number of EPUs 34 with a different relative number of fixed EPUs 56 and tilting EPUs 54 (or indeed the fixed EPUs 56 or the tilting EPUs 54 may be omitted entirely), and a different arrangement relative to the left / right plane 22 and the fore / aft plane 22. For example, the eVTOL aircraft 10 may comprise six tilting EPUs 54.

[0232] In the above examples described with reference to Figures 1 to 17, each busbar 32 is connected to two EPUs 34. In other examples, each busbar 32 may be connected to more than two EPUs 34. In the above example, each inverter 82 is connected to a single busbar 32. In other examples, each inverter 82 may be connected to multiple busbars 32. In the above examples, each motor 60 comprises two windings 68 and each EPU 34 comprises a dual lane inverter apparatus 64 comprising two inverters 82. In other examples, each motor 60 may comprise two or more windings 68. Accordingly, each EPU 34 may comprises a multi-lane inverter apparatus 64 comprising two or more inverters 82. In this case, the voltage controller 83 may receive a plurality of measurements. In the above example, each dual lane inverter apparatus 64 comprises a pair of inverters 82. In other examples, each dual lane inverter apparatus 64 may comprise more than two inverters 82 and may be referred to as a multi lane inverter apparatus.

[0233] In the above examples described with reference to Figures 1 to 17, the measurements received by each voltage controller 83 are voltage measurements which are indicative of the discharge rate of the respective battery 30. Additionally, the voltage controllers 83 employ the proportional regime and regenerative regime to reduce the difference in the discharge rates of the batteries 30. The discharge rate is one example of a power delivery property of the batteries 30. In other examples, other power delivery properties such as the state of charge or temperature of the batteries 30 may be used instead. In these examples, the measurements received will be indicated of the power delivery properties, and the voltage controllers 83 will employ the proportional regime and the regenerative regimes to reduce the difference in the power delivery properties.

[0234] In the above examples described with reference to Figures 7 to 17, when operating in the proportional regime, the voltage controllers 83 determine a new ratio of the levels of power of each inverter 82 by varying the previous ratio of the levels of power proportionally to the difference in the voltages indicated by the pair of measurements. In other examples, other kinds of proportional regime may be employed where the levels of power are not exactly proportional to the difference in the voltages. For example, the new ratio may be determined based on a proportional, an integral, and a differential term calculated based on the difference in the voltages indicated by the pair of measurements. In another example, the previous ratio may be varied based on the voltages indicated by the pair of measurements. For example, the ratio could be varied such that the level of power of the inverter with the higher voltage is increased and the level of power of the inverter with the lower voltage is reduced, without necessarily being proportional to the difference in voltages. As mentioned above, the measurements received may be indicated of the power delivery properties. In these examples, the proportional regime may comprise varying the previous ratio of the levels of power proportionally to the difference in the power delivery properties indicated by the pair of measurements.

[0235] In the above examples described with reference to Figures 1 to 17, each EPU 34 comprises a respective voltage controller 83. In other examples, the respective voltage controllers 83 may be replaced with a single centralised control system which performs the functions of the respective voltage controllers 83.

Claims

Claims1. An electric vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, and a plurality of electric propulsion units, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements; and control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements.

2. The electric vertical take-off and landing aircraft of any one of the preceding claims, wherein the power delivery property is a discharge rate.

3. The electric vertical take-off and landing aircraft of claim 1 or claim 2, wherein each measurement of the plurality of measurements is a measurement of a voltage of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements.

4. The electric vertical take-off and landing aircraft of any one of the preceding claims, wherein each power control element of the plurality of power control elements is an inverter; and the electric motor is an alternating current electric motor.

5. The electric vertical take-off and landing aircraft of claim 4, wherein the electric motor is a three-phase alternating current electric motor.

6. The electric vertical take-off and landing aircraft of any one of the preceding claims, wherein: the controller is configured to control the plurality of power control elements according to a regenerative regime when one of the respective batteries of the plurality of batteries connected to each power control element of the plurality of power control elements has a different power delivery property than another of the respective batteries of the plurality of batteries connected to each power control element of the plurality of power control elements; and the regenerative regime comprises the winding of the plurality of windings which is connected to the respective battery with the different power delivery property: regeneratively harvesting energy from the electric motor; and suppling the regeneratively harvested energy to the power distribution assembly connected to the winding which is regeneratively harvesting the energy in order to reduce a difference in power delivery properties indicated by the received plurality of measurements.

7. The electric vertical take-off and landing aircraft of claim 6, wherein: the controller is configured to switch between controlling the plurality of power control elements according to the regenerative regime and controlling the plurality of power control elements according to a proportional regime; and the proportional regime comprises varying the level of power delivered to each winding of the plurality of windings from the respective battery connected to each power control element in proportion to the difference in power delivery properties indicated by the received plurality of measurements.

8. The electric vertical take-off and landing aircraft of claim 7, wherein the controller is configured to switch from controlling the plurality of power control elements according to the proportional regime to controlling the plurality of power control elements according to the regenerative regime based on a maximum power rating of the power control element connected to the other of the respective batteries of the plurality of batteries being greater than a total level of power delivered to the plurality of windings when the controller controls the plurality of power control elements according to the proportional regime.

9. The electric vertical take-off and landing aircraft of claim 7 or claim 8, wherein: the power delivery property is a discharge rate; the different power delivery property is a higher discharge rate; the controller is configured to switch from controlling the plurality of power control elements according to the proportional regime to controlling the plurality of power control elements according to the regenerative regime based on a discharge rate of the other of the respective batteries of the plurality of batteries connected to each power control element of the plurality of power control elements being less than or equal to a higher discharge rate threshold.

10. The electric vertical take-off and landing aircraft of any one of the preceding claims, wherein: the electric motor has a maximum power rating; and each power control element of the plurality of power control elements has a maximum power rating of no less than 60% and no greater than 85% of the maximum power rating of the electric motor.

11. The electric vertical take-off and landing aircraft of any one of the preceding claims, wherein: the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; anda fore / aft plane which extends parallel to the pitch axis and the yaw axis, perpendicular to the roll axis, and through the centre of gravity; the electric vertical take-off and landing aircraft comprises a flight control computer configured to, in response to detecting a failure of one of the electric propulsion units of the plurality of electric propulsion units, send a power command to the controller of at least one of the electric propulsion units; and the controller is configured to, in response to receiving the power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command and to reduce a difference in power delivery property indicated by the received plurality of measurements.

12. The electric vertical take-off and landing aircraft of claim 11, wherein the flight control computer is configured to: in response to detecting a failure of an electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the left / right plane, send the power command to another electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane, and wherein the power command is a command to increase the power of the motor; and / or in response to detecting a failure of a further electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the fore / aft plane, send the power command to a different electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the fore / aft plane, and wherein the power command is a command to increase the power of the electric motor.

13. The electric vertical take-off and landing aircraft of claim 11 or claim 12, wherein the flight control computer is configured to: in response to detecting a failure of an electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane and the first side of the fore / aft plane, send the power command to another electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a second side of the left / right plane and a second side of the fore / aft plane, and wherein the power command is a command to decrease the power of the electric motor.

14. The electric vertical take-off and landing aircraft of claim 13, wherein for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units are: mounted to the airframe on the same side of the left / right plane and / or the same side of the fore / aft plane; and the only electric propulsion units to which the respective power distribution assembly is connected.

15. The electric vertical take-off and landing aircraft of any one of the preceding claim, wherein the controller is configured to, in response to a failure of one of the power control elements of the plurality of power control elements, control a non-failed power control element of the plurality of power control elements to increase the level of power that is delivered to the respective winding.

16. The electric vertical take-off and landing aircraft of any one of the preceding claim, wherein the controller is configured to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the power control elements of the plurality of power control elements, control a non-affected power control element of the plurality of power control elements which is not connected to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding.

17. The electric vertical take-off and landing aircraft of claim 16, wherein the level of power that is delivered to the respective winding connected to the non-affected power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

18. The electric vertical take-off and landing aircraft of any one of the preceding claims, wherein: the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; anda left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units to which a respective power distribution assembly of the majority of the plurality of power distribution assemblies is connected are mounted to the airframe on a same sides of the left / right plane.

19. The electric vertical take-off and landing aircraft of any one of the preceding claim, wherein: the multiple electric propulsion units to which a respective power distribution assembly of the plurality of power distribution assemblies is connected comprise a tilting electric propulsion unit and a fixed electric propulsion unit; the electric vertical take-off and landing aircraft comprises a rotor, a tilting mechanism, and a further rotor; the tilting electric propulsion unit is connected to the rotor; the tilting mechanism is configured to tilt the rotor relative to the airframe between a vertical flight mode in which the rotor delivers thrust which is parallel to the yaw axis and a convention flight mode in which the rotor delivers thrust which is parallel to the roll axis of the aircraft in use; and the fixed electric propulsion unit is connected to the further rotor which is fixed relative to airframe such that, in use, the further rotor delivers thrust which is parallel to the roll axis of the aircraft in use only.

20. The electric vertical take-off and landing aircraft of any one of the preceding claim, wherein each power distribution assembly of the plurality of power distribution assemblies comprises a busbar.

21. The electric vertical take-off and landing aircraft of any one of claims 1 to 5, wherein each power control element of the plurality of power control elements is unidirectional.

22. The electric vertical take-off and landing aircraft of any one of the preceding claims, wherein: the controller is configured to control the plurality of power control elements, when power is deliverable from the respective batteries of the plurality of batteries to the respectivewindings, to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

23. The electric vertical take-off and landing aircraft of claim 22, wherein: the higher threshold is no less than 90% of the total level of power delivered to the respective windings; and the lower threshold is no greater than 10% of the total level of power delivered to the respective windings.

24. An electric propulsion unit for an electric vertical take-off and landing aircraft, the electric propulsion unit comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; andcontrol the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

25. A multi -lane inverter apparatus for an electric propulsion unit of an electric vertical take-off and landing aircraft; the multi-lane inverter apparatus comprising: a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of winding connections, each winding connection of the plurality of winding connections is connectable to a respective winding of a plurality of windings of an alternating current electric motor of the electric propulsion unit; a plurality of inverters, each inverter of the plurality of inverters is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding connection of the plurality of winding connections such that each inverter can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of the respective battery of the plurality of batteries connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; and control the plurality of inverters to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements.

26. An electric vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electric propulsion units, and a control system, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distributionassembly is connected to multiple electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding, wherein the control system is configured, for an electric propulsion unit of the plurality of electric propulsion units, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery connected to a respective power control element of the electric propulsion unit; and control a power control element of the electric propulsion unit according to a regenerative regime when the respective battery to which the power control element is connected has a higher discharge rate than the, or each, other respective battery connected to the, or each, other power control element of the electric propulsion unit, wherein the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

27. The electric vertical take-off and landing aircraft of claim 26, wherein: the control system is configured, for the electric propulsion unit, to switch between controlling the power control element according to the regenerative regime and controlling the power control element according to a proportional regime; and the proportional regime comprises the power control element varying the level of power delivered to the respective winding connected to the power control element from the respective battery connected to the power control element in proportion to the difference in discharge rates indicated by the received plurality of measurements.

28. The electric vertical take-off and landing aircraft of claim 27, wherein the control system is configured, for the electric propulsion unit, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based on a maximum power rating of another power control element of the electric propulsion unit being greater than a total level of power delivered to the plurality of windings when the control system controls the power control element according to the proportional regime.

29. The electric vertical take-off and landing aircraft of claim 27 or claim 28, wherein: the control system is configured, for the electric propulsion unit, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based on a discharge rate of the respective battery connected to another power control element of the electric propulsion unit being less than or equal to a higher discharge rate threshold.

30. The electrical vertical take-off and landing aircraft of any one of claims 26 to 29, wherein: the electrical vertical take-off and landing aircraft comprises a flight control computer configured to send respective power commands for at least two electric propulsion units of the plurality of electric propulsion units to the control system; the control system is configured, for each electric propulsion unit of the at least two electric propulsion units to, in response to receiving the respective power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the respective power command and to reduce a difference in discharge rates indicated by the received plurality of measurements; and at least two of the respective power commands are different from one another.

31. The electric vertical take-off and landing aircraft of any one of claims 26 to 30, wherein each measurement of the plurality of measurements is a measurement of a voltage of a respective battery connected to a respective power control element of the electric propulsion unit.

32. The electric vertical take-off and landing aircraft of any one of claims 26 to 31, wherein each power control element of the plurality of power control elements is an inverter; and the electric motor is an alternating current electric motor.

33. The electric vertical take-off and landing aircraft of claim 32, wherein the electric motor is a three-phase alternating current electric motor.

34. The electric vertical take-off and landing aircraft of any one of claims 26 to 33, wherein: the electric motor has a maximum power rating; and each power control element of the plurality of power control elements has a maximum power rating of no less than 60% and no greater than 85% of the maximum power rating of the electric motor.

35. The electric vertical take-off and landing aircraft of any one of claims 26 to 34, wherein: the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and a fore / aft plane which extends parallel to the pitch axis and the yaw axis, perpendicular to the roll axis, and through the centre of gravity; the electric vertical take-off and landing aircraft comprises a flight control computer configured to, in response to detecting a failure of one of the electric propulsion units of the plurality of electric propulsion units, send a power command for the electric propulsion unit to the control system; and the control system is configured, for the electric propulsion unit, to, in response to receiving the power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command and to reduce a difference in discharge rate indicated by the received plurality of measurements.

36. The electric vertical take-off and landing aircraft of claim 35, wherein the flight control computer is configured such that:in response to detecting a failure of an additional electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the left / right plane, the electric propulsion unit is an alternate electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane, and wherein the power command is a command to increase the power of the electric motor; and / or in response to detecting a failure of a further electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the fore / aft plane, the electric propulsion unit is a different electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the fore / aft plane, and wherein the power command is a command to increase the power of the electric motor.

37. The electric vertical take-off and landing aircraft of claim 35 or claim 36, wherein the flight control computer is configured such that: in response to detecting a failure of a failed electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane and the first side of the fore / aft plane, the electric propulsion unit is a further electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a second side of the left / right plane and a second side of the fore / aft plane, and wherein the power command is a command to decrease the power of the electric motor.

38. The electric vertical take-off and landing aircraft of claim 37, wherein for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units are: mounted to the airframe on the same side of the left / right plane and / or the same side of the fore / aft plane; and the only electric propulsion units to which the respective power distribution assembly is connected.

39. The electric vertical take-off and landing aircraft of any one claims 26 to 38, wherein the control system is configured, for the electric propulsion unit, to, in response to a failure of one of the power control elements of the plurality of power control elements, control a nonfailed power control element of the plurality of power control elements to increase the level of power that is delivered to the respective winding.

40. The electric vertical take-off and landing aircraft of any one of claims 26 to 39, wherein the control system is configured, for the electric propulsion unit, to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the power control elements of the plurality of power control elements, control a non-affected power control element of the plurality of power control elements which is not connected to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding.

41. The electric vertical take-off and landing aircraft of claim 40, wherein the level of power that is delivered to the respective winding connected to the non-affected power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

42. The electric vertical take-off and landing aircraft of any one of claims 26 to 41, wherein: the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; and a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units to which a respective power distribution assembly of the majority of the plurality of power distribution assemblies is connected are mounted to the airframe on a same sides of the left / right plane.

43. The electric vertical take-off and landing aircraft of any one of claims 26 to 42, wherein: the multiple electric propulsion units to which a respective power distribution assembly of the plurality of power distribution assemblies is connected comprise a tilting electric propulsion unit and a fixed electric propulsion unit; the electric vertical take-off and landing aircraft comprises a rotor, a tilting mechanism, and a further rotor; the tilting electric propulsion unit is connected to the rotor;the tilting mechanism is configured to tilt the rotor relative to the airframe between a vertical flight mode in which the rotor delivers thrust which is parallel to the yaw axis and a convention flight mode in which the rotor delivers thrust which is parallel to the roll axis of the aircraft in use; and the fixed electric propulsion unit is connected to the further rotor which is fixed relative to airframe such that, in use, the further rotor delivers thrust which is parallel to the roll axis of the aircraft in use only.

44. The electric vertical take-off and landing aircraft of any one of claims 26 to 43, wherein each power distribution assembly of the plurality of power distribution assemblies comprises a busbar.

45. The electric vertical take-off and landing aircraft of any one of claims 26 to 44, wherein: the control system is configured, for the electric propulsion unit, to control the plurality of power control elements of the electric propulsion unit, when power is deliverable from the respective batteries to the respective windings of the electric propulsion unit, to vary the level of power delivered to the respective windings in order to reduce a difference in the discharge rates indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

46. The electric vertical take-off and landing aircraft of claim 45, wherein: the higher threshold is no less than 90% of the total level of power delivered to the respective windings; and the lower threshold is no greater than 10% of the total level of power delivered to the respective windings.

47. The electric vertical take-off and landing aircraft of any one of claims 26 to 46, wherein: the control system comprises a plurality of controllers; and each electric propulsion unit of the plurality of electric propulsion units comprises a respective controller of the plurality of controllers; and each controller is configured, for a respective electrical propulsion unit to:control a power control element of the respective electric propulsion unit according to a regenerative regime when the respective battery to which the power control element is connected has a higher discharge rate than the, or each, other respective battery connected to the, or each, other power control element of the respective electric propulsion unit, and the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

48. A control system for an electric vertical take-off and landing aircraft, the electrical vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electric propulsion units, and a control system, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings, wherein the control system is configured, for an electric propulsion unit of the plurality of electric propulsion units, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery connected to a respective power control element of the plurality of the electric propulsion unit; andcontrol a power control element of the electric propulsion unit according to a regenerative regime when the respective battery to which the power control element is connected has a higher discharge rate than the, or each, other respective battery of the plurality of batteries connected to the, or each, other power control element of the electrical propulsion unit, wherein the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

49. An electric propulsion unit for an electric vertical take-off and landing aircraft, the electric propulsion unit comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections in use; and control a power control element of the plurality of power control elements according to a regenerative regime when the respective battery to which the power control element is connected in use has a higher discharge rate than the, or each, other respective battery connected in use to the, or each, other respective power control element of the plurality of power control elements,wherein the regenerative regime comprises the power control element causing the winding that is connected in use to the respective battery with the higher discharge rate to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in use in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

50. A multi-lane inverter apparatus for an electric propulsion unit of an electric vertical take-off and landing aircraft; the multi-lane inverter apparatus comprising: a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of winding connections, each winding connection of the plurality of winding connections is connectable to a respective winding of a plurality of windings of an alternating current electric motor of the electric propulsion unit; a plurality of inverters, each inverter of the plurality of inverters is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respective winding connection of the plurality of winding connections such that each inverter can control a level of power that is delivered from the respective battery to the respective winding of the plurality of windings in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a discharge rate of a respective battery of the plurality of batteries connected to each power distribution assembly connection of the plurality of power distribution assembly connections in use; and control an inverter of the plurality of inverters according to a regenerative regime when the respective battery to which the inverter is connected in use has a higher discharge rate than the, or each, other respective battery connected in use to the, or each, respective inverter of the plurality of inverters, wherein the regenerative regime comprises the inverter causing the winding that is connected in use to the respective battery with the higher discharge rate to regeneratively harvest energy from the alternating current electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in use in order to reduce a difference in discharge rate indicated by the received plurality of measurements.

51. An electric vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, a plurality of electrical propulsion units, and a control system, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electrical motor comprising a plurality of windings; and a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery of the plurality of batteries to the respective winding of the plurality of windings, and wherein, the control system is configured, for an electric propulsion unit of the plurality of electric propulsion units, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power control element of the electric propulsion unit; and when power is deliverable from the respective batteries to the respective windings of the electric propulsion unit, control the plurality of power control elements of the electric propulsion unit to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

52. The electric vertical take-off and landing aircraft of claim 51, wherein: the higher threshold is no less than 90% of the total level of power delivered to the respective windings; and / or the lower threshold is no greater than 10% of the total level of power delivered to the respective windings.

53. The electric vertical take-off and landing aircraft of claim 51 or 52, wherein the power delivery property is a discharge rate.

54. The electric vertical take-off and landing aircraft of any one of claims 51 to 53, wherein each measurement of the plurality of measurements is a measurement of a voltage of the respective battery of the plurality of batteries connected to each power control element of the plurality of power control elements.

55. The electric vertical take-off and landing aircraft of any one of claims 51 to 54, wherein each power control element of the plurality of power control elements is an inverter; and the electric motor is an alternating current electric motor.

56. The electric vertical take-off and landing aircraft of claim 55, wherein the electric motor is a three-phase alternating current electric motor.

57. The electric vertical take-off and landing aircraft of any one of claimsa 51 to 56, wherein: the control system is configured, for the electric propulsion unit, to control a power control elements of the electric propulsion unit according to a regenerative regime when the respective battery to which the power control element is connected has a different power delivery property the, or each, other respective battery of the plurality of batteries connected to the, or each, other power control element of the electric propulsion unit; and the regenerative regime comprises the power control element causing the winding that is connected to the respective battery with the different power delivery property to regeneratively harvest energy from the electric motor and supply the regeneratively harvested energy to the power distribution assembly to which it is connected in order to reduce a difference in power delivery properties indicated by the received plurality of measurements.

58. The electric vertical take-off and landing aircraft of claim 57, wherein: the control system is configured, for the electric propulsion unit, to switch between controlling the power control element according to the regenerative regime and controlling the power control elements according to a proportional regime; and the proportional regime comprises the power control element varying the level of power delivered to the winding connected to the power control element from the respective battery connected to the power control element in proportion to the difference in power delivery properties indicated by the received plurality of measurements.

59. The electric vertical take-off and landing aircraft of claim 58, wherein the control system is configured, for the electric propulsion unit, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based on a maximum power rating of another power control element of the electric propulsion unit being greater than a total level of power delivered to the plurality of windings when the control system controls the power control element according to the proportional regime.

60. The electric vertical take-off and landing aircraft of claim 58 or claim 59, wherein: the power delivery property is a discharge rate; the different power delivery property is a higher discharge rate; the control system is configured, for the electric propulsion unit, to switch from controlling the power control element according to the proportional regime to controlling the power control element according to the regenerative regime based on a discharge rate of the respective battery connected to another power control element of the electric propulsion unit being less than or equal to a higher discharge rate threshold.

61. The electric vertical take-off and landing aircraft of any one of claims 51 to 60, wherein: the electric motor has a maximum power rating; and each power control element of the plurality of power control elements has a maximum power rating of no less than 60% and no greater than 85% of the maximum power rating of the electric motor.

62. The electric vertical take-off and landing aircraft of any one of claims 51 to 61, wherein:the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and a fore / aft plane which extends parallel to the pitch axis and the yaw axis, perpendicular to the roll axis, and through the centre of gravity; the electric vertical take-off and landing aircraft comprises a flight control computer configured to, in response to detecting a failure of one of the electric propulsion units of the plurality of electric propulsion units, send a power command for the electric propulsion unit to the control system; and the control system is configured, for the electric propulsion unit, to, in response to receiving the power command, control the plurality of power control elements to vary the level of power delivered to the respective windings based on the power command and to reduce a difference in power delivery property indicated by the received plurality of measurements.

63. The electric vertical take-off and landing aircraft of claim 62, wherein the flight control computer is configured such that: in response to detecting a failure of an additional electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the left / right plane, the electric propulsion unit is an alternate electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane, and wherein the power command is a command to increase the power of the electric motor; and / or in response to detecting a failure of a further electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a first side of the fore / aft plane, the electric propulsion unit is a different electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the fore / aft plane, and wherein the power command is a command to increase the power of the electric motor.

64. The electric vertical take-off and landing aircraft of claim 62 or claim 63, wherein the flight control computer is configured such that:in response to detecting a failure of a failed electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on the first side of the left / right plane and the first side of the fore / aft plane, the electric propulsion unit is a further electric propulsion unit of the plurality of electric propulsion units which is mounted to the airframe on a second side of the left / right plane and a second side of the fore / aft plane, and wherein the power command is a command to decrease the power of the electric motor.

65. The electric vertical take-off and landing aircraft of claim 64, wherein for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units are: mounted to the airframe on the same side of the left / right plane and / or the same side of the fore / aft plane; and the only electric propulsion units to which the respective power distribution assembly is connected.

66. The electric vertical take-off and landing aircraft of any one of claims 51 to 65, wherein the control system is configured, for the electric propulsion unit, to, in response to a failure of one of the power control elements of the plurality of power control elements, control a nonfailed power control element of the plurality of power control elements to increase the level of power that is delivered to the respective winding.

67. The electric vertical take-off and landing aircraft of any one of claims 51 to 66, wherein the control system is configured, for the electric propulsion unit, to, in response to a failure of the respective battery and / or the respective power distribution assembly connected to one of the power control elements of the plurality of power control elements, control a non-affected power control element of the plurality of power control elements which is not connected to the failed power distribution assembly and / or the failed battery to increase the level of power that is delivered to the respective winding.

68. The electric vertical take-off and landing aircraft of claim 67, wherein the level of power that is delivered to the respective winding connected to the non-affected power control element is increased to greater than or equal to 90% of a total level of power that was delivered to the respective windings immediately before the failure of the respective battery and / or the respective power distribution assembly.

69. The electric vertical take-off and landing aircraft of any one of claims 51 to 68, wherein: the electric vertical take-off and landing aircraft comprises: an airframe; a centre of gravity; a pitch axis, a roll axis, and a yaw axis; and a left / right plane which extends parallel to the roll axis, perpendicular to the pitch axis and the yaw axis, and through the centre of gravity; and for a majority of the power distribution assemblies of the plurality of power distribution assemblies, the multiple electric propulsion units to which a respective power distribution assembly of the majority of the plurality of power distribution assemblies is connected are mounted to the airframe on a same sides of the left / right plane.

70. The electric vertical take-off and landing aircraft of any one of claims 51 to 69, wherein: the multiple electric propulsion units to which a respective power distribution assembly of the plurality of power distribution assemblies is connected comprise a tilting electric propulsion unit and a fixed electric propulsion unit; the electric vertical take-off and landing aircraft comprises a rotor, a tilting mechanism, and a further rotor; the tilting electric propulsion unit is connected to the rotor; the tilting mechanism is configured to tilt the rotor relative to the airframe between a vertical flight mode in which the rotor delivers thrust which is parallel to the yaw axis and a convention flight mode in which the rotor delivers thrust which is parallel to the roll axis of the aircraft in use; and the fixed electric propulsion unit is connected to the further rotor which is fixed relative to airframe such that, in use, the further rotor delivers thrust which is parallel to the roll axis of the aircraft in use only.

71. The electric vertical take-off and landing aircraft of any one of claims 51 to 70, wherein each power distribution assembly of the plurality of power distribution assemblies comprises a busbar.

72. The electric vertical take-off and landing aircraft of any one of claims 51 to 71, wherein each power control element of the plurality of power control elements is unidirectional.

73. The electric vertical take-off and landing aircraft of any one of claims 51 to 72, wherein: the control system comprises a plurality of controllers; each electric propulsion unit of the plurality of electric propulsion units comprises a respective controller of the plurality of controllers; and each controller is configured, for a respective electric propulsion unit, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power control element of the electric propulsion unit; and when power is deliverable from the respective batteries of the plurality of batteries to the respective windings of the electric propulsion unit, control the plurality of power control elements of the electric propulsion unit to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

74. A control system for an electric vertical take-off and landing aircraft, the electrical vertical take-off and landing aircraft comprising: a plurality of batteries, a plurality of separate power distribution assemblies, and a plurality of electric propulsion units, wherein: each battery of the plurality of batteries is connected to a respective power distribution assembly of the plurality of power distribution assemblies, and each power distribution assembly is connected to multiple electric propulsion units of the plurality of electric propulsion units; each electric propulsion unit of the plurality of electric propulsion units comprises: an electric motor comprising a plurality of windings; anda plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective battery of the plurality of batteries via a respective power distribution assembly of the plurality of power distribution assemblies and to a respective winding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding, and wherein the control system configured, for an electric propulsion unit of the plurality of electric propulsion units, to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power control element of the electric propulsion unit; and when power is deliverable from the respective batteries to the respective windings of the electric propulsion unit, control the plurality of power control elements of the electric propulsion unit to vary the level of power delivered to the respective windings in order to reduce a difference in the power delivery properties indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no less than a higher threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no greater than a lower threshold of the total level of power delivered to the respective windings.

75. An electric propulsion unit for an electric vertical take-off and landing aircraft, the electric propulsion unit comprising: an electric motor comprising a plurality of windings; a plurality of power distribution assembly connections, each power distribution assembly connection of the plurality of power distribution assembly connections is connectable to a respective battery of the electric vertical take-off and landing aircraft via a power distribution assembly of the electric vertical take-off and landing aircraft; a plurality of power control elements, each power control element of the plurality of power control elements is connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections and to a respectivewinding of the plurality of windings such that each power control element can control a level of power that is delivered from the respective battery to the respective winding in use; and a controller configured to: receive a plurality of measurements, each measurement of the plurality of measurements is indicative of a power delivery property of a respective battery connected to a respective power distribution assembly connection of the plurality of power distribution assembly connections in use; and when power is deliverable from the respective batteries to the respective windings, control the plurality of power control elements to vary the level of power delivered to the respective windings in order to reduce a difference in power delivery property indicated by the received plurality of measurements, and such that: the level of power delivered to one of the respective windings is no greater than a lower threshold of a total level of power delivered to the respective windings; and the level of power delivered to another of the respective windings is no less than a higher threshold of the total level of power delivered to the respective windings.

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