A delivery drone

Abstract

The present application provides an unmanned aerial vehicle for the delivery of product comprising a plurality of drive units. Each drive unit drives an associated aero-rotor. Each drive unit comprises at least two motors driving the same aero-rotor.

Description

[0001] A Delivery Drone

[0002] Field of the Application

[0003] The present application relates to aircraft and in particular to a drone for delivering products to a customer.

[0004] Background of the Application

[0005] A drone is an unmanned aerial vehicle (UAV). A particular application of drones is product delivery. Product delivery imposes a number of restrictions on the design of a drone, including for example, that the drone is intended for use over populated areas.

[0006] The present application is directed to an improved drone for use in product delivery.

[0007] Summary

[0008] The present application provides, in order to address one or more issues of reliability, safety, efficiency and recovery in the event of a failure, a dual motor multirotor propulsion system for an aircraft.

[0009] Accordingly, a first aspect provides an unmanned aerial vehicle. The unmanned aerial vehicle comprises an aircraft body. A plurality of aero-rotors provided about the aircraft body, with each aero-rotor being arranged to rotate to provide lift for the aircraft. A payload delivery system may be provided for delivering a payload at a desired delivery location. The unmanned aerial vehicle has a plurality of drive units, each drive unit being configured to rotate an associated aero-rotor, with each drive unit comprising a plurality of electric motors. Each individual electric motor of the plurality of electric motors alone or in combination with the other electric motor(s) of the plurality is operable to rotate the associated aero-rotor.

[0010] To power the unmanned aerial vehicle, at least one energy supply system is provided which powers the drive units. A flight control system is provided for controlling the aircraft, the flight control system being operable to operate the drive units to manoeuvre the unmanned aircraft to a desired delivery location.

[0011] Each drive unit may be designed to provide a maximum operational torque, MOT, and where each motor of the drive unit is configured to provide a maximum operational torque of a M0T / (N-1), where N is the number of electric motors in a drive unit and a is a factor determine a reduced maximum operational torque in the event of a failure.

[0012] Each motor has a motor rotor rotating about a rotor shaft. The rotor shafts of each motor may be connected together. The rotor shafts may be co-axial.

[0013] The motors of each drive unit may share a common rotor shaft. The motors of each drive unit may be co-housed and provided as a single unit. In this arrangement, each motor may comprise a separately excitable motor stator with the motors sharing a common motor rotor. In this arrangement, the motor stators may be wound together as different segments on a segmented stator.

[0014] Each drive unit may be directly connected to an associated aero-rotor without the use of a gearbox or the rotor shafts of each motor may be connected through a fly or gear wheel to the aero-rotor.

[0015] Where, each drive unit has a first motor and a second motor, the first and second motors in combination, may be configured to provide a maximum operational torque, to drive the aero-rotor and wherein each of the first and second motors is individually capable of delivering a proportion of the maximum operational torque, the proportion being greater than 70%.

[0016] A separate electronic speed controller may be associated with each motor of each drive unit. Each electronic speed controller may be configured in response to one or more signals from the flight control system to control the power delivered from the at least one electrical power source to the motor associated with that electronic speed controller.

[0017] In one variation, a single electronic speed controller may be provided for each drive unit, the single electronic speed controller being associated with all of the motors in the drive unit, the electronic speed controller being configured in response to one or more signals from the flight control system to control the power delivered from the at least one electrical power source to the individual motors of the drive unit.

[0018] Each drive unit may comprise a housing. The one or more electronic speed controllers for a drive unit may be integrated with the motors in the drive unit.

[0019] The housing may have a releasable connection which co-operates with corresponding features on the aircraft body allowing for the drive unit housing to be removed / replaced as required. The releasable connection is suitably a quick connect-disconnect connection.

[0020] The aircraft may have a plurality of energy supply systems. In such a configuration, each of the plurality of energy supply systems may uniquely supply energy to a motor in each of the drive units so as to ensure that in the event of a failure of one energy supply system, only those motors connected to that energy supply system will fail. The remaining motors can continue to operate being supplied by a remaining energy supply system.

[0021] An energy supply system may supply electrical energy to each drive unit. The energy supply system may have a plurality of cells arranged as multiple series strings in parallel. In this arrangement, individual strings of the series may be electrically disconnected in the event of a failure.

[0022] Similarly, the energy supply system may have a plurality of cells connected as multiple parallel strings arranged in series. In this arrangement, individual cells can be disconnected in the event of a failure. The energy supply system may have cells of different battery chemistry arranged in a multi-battery chemistry pack providing a single output to each drive unit.

[0023] In such an arrangement, a first battery chemistry may be selected to provide a first cell with a first C-rate and having a first energy capacity and a second battery chemistry may be selected to provide a second cell with a second C-rate and having a second energy capacity, wherein the first C-rate is lower than the second C-rate and the first energy capacity is higher than the second energy capacity.

[0024] The energy supply system may comprise cells of different battery chemistry arranged in a multi-battery chemistry pack providing at least two outputs to each drive unit with each output powering a different motor of the drive unit.

[0025] In such a configuration, the motors of the drive unit may have different operating characteristics with each motor being selected to match the battery chemistry providing the output to which the motor is connected.

[0026] In flight, one of the first and second motors may be selected to be an active motor which in flight is powered to drive the aero-rotor with the remaining motor selected to be a passive motor which is unpowered. In this configuration, upon the failure of active motor, the passive motor is selected to assume the role of the active motor.

[0027] Alternatively, the first and second motors, in flight, share the drive load substantially equally. In this alternative, the ratio of power delivered by each motor may varied dynamically during flight. This variation may be based on current flight conditions.

[0028] Similarly, the ratio of power may be varied to account for thermal differences between motors or semiconductor devices controlling the motors.

[0029] The payload delivery system for delivering a payload at a desired delivery location may be configured to collect the payload from a sender’s location. In such a configuration, it may be configured to unwind a tether whilst hovering to a waiting dispatcher below.

[0030] In another aspect, the present application provides an unmanned aerial vehicle having an aircraft body with a plurality of rotors provided about the aircraft body, each rotor being arranged to rotate to provide lift for the aircraft.

[0031] A plurality of drive units are provided with each drive unit being configured to rotate an associated aero-rotor with each drive unit comprising a motor having a rotor and at least two stators wound together on the same former.

[0032] At least one energy supply system providing power to the drive units. A flight control system is provided to control the aircraft, the flight control system being operable to operate the drive units to manoeuvre the unmanned aircraft to a desired delivery location.

[0033] An electronic speed controller may be associated with each stator, with each electronic speed controller being configured in response to a signal from the flight control system to control the power delivered from the at least one electrical power source to the associated stator. There may be a plurality of energy supply systems, each of the plurality of energy supply systems uniquely suppling energy to a stator in each of the drive units. This ensures that in the event of a failure of one energy supply system, that only those stators connected to that energy supply system will fail and the remaining stators can continue to operate being supplied by the remaining energy supply system.

[0034] One of the first and second stators may be selected to be an active stator which in flight is powered with the remaining stator selected to be a passive motor which is unpowered. Upon the failure of active stator, the passive stator is selected to assume the role of the active stator.

[0035] Alternatively, the first and second stators, in flight, share the drive load substantially equally.

[0036] A further aspect of the application provides a drone, comprising four aero-rotors and four drive units. Each drive unit is operable to rotate an associated aero-rotor. In this aspect, each drive unit comprises two stators and at least one rotor. In this further aspect, eight electronic speed controllers are provided with each electronic speed controller converting a DC voltage to an AC voltage and providing the AC voltage to an associated stator. In any given drive unit, the stators may both be active or a single stator may be active. Alternatively, in flight both stators may be active. In this case, the power split between the stators may be changed dynamically.

[0037] Another aspect relates to an electric propulsion architecture for a dual motor multirotor drone for drone delivery service. The drone suitably comprises at least 8 electric motors, with at least 8 electronic speed controllers. Each electronic speed controller converts a DC voltage to an AC voltage to control the current, torque, and / or speed of the electric motors. The architecture may have a single or redundant power supply systems (including ESS, connectors, fuses, busbars, cabling etc.). Suitably the architecture is directed to having at least 4 aero-rotors.

[0038] For each aero-rotor, there are two electric motors wherein either both can be active or a single motor can be active and the power split between both can be dynamically changed. In the event of a failure, one motor will still remain functional. Each aerorotor may be mechanically coupled to both electric motors such that during operation both, or either, can drive the aero-rotor.

[0039] Other systems, devices, methods, features and advantages of the subject matter described herein will be or will become apparent to one with skill in the art upon examination of the following figures and detailed description. It is intended that all such additional systems, methods, features and advantages be included within this description, be within the scope of the subject matter described herein, and be protected by the accompanying claims. In no way should the features of the example embodiments be construed as limiting the appended claims, absent express recitation of those features in the claims. Brief Description of Drawings

[0040] The accompanying drawings, which are included as part of the present specification, illustrate the presently example embodiments and, together with the general description given above and the detailed description of the example embodiments given below, serve to explain and teach the principles of the present invention.

[0041] Figure 1 is an illustration of a drone according to the first aspect of the present application;

[0042] Figure 2 is an illustration of a method of flight control using motor control for a drone of the type shown generally in Figure 1 ;

[0043] Figure 3 is an exemplary system view for the drone of Figure 1 ;

[0044] Figure 4 is an illustration of a payload delivery system for use with the drone of Figure 1 ;

[0045] Figure 5 is an illustration of an exemplary first rotor-motor assembly in which two motors are employed to drive a common aero-rotor;

[0046] Figure 6 is an illustration of an exemplary second rotor-motor assembly in which two motors are employed to drive a common aero-rotor through a flywheel;

[0047] Figure 7 is an illustration of a motor mount that allows for rapid replacement of a motor assembly and associated ESCs.

[0048] Figure 8 is an illustration of a control system suitable for control of the arrangements of Figures 5 to 7 with a single energy storage system;

[0049] Figure 9 illustrates an exemplary configuration of 12 battery cells in series with 4 parallel strings (13S4P) for an energy storage system for a drone;

[0050] Figure 10 illustrates another exemplary configuration of 4 cells in parallel with 13 of these in series (4P13S) for an energy storage system for a drone;

[0051] Figure 11 illustrates an example of power sharing based on the control system shown in Figure 8 during normal operation;

[0052] Figure 12 illustrates an example of power sharing for the arrangement of Figure 8 exhibiting a motor fault;

[0053] Figure 13 is an illustration of a control system suitable for control of the arrangements of Figures 5 to 7 with a dual power supply system;

[0054] Figure 14 illustrates the effect on the arrangement of Figure 13 when there is energy supply system fault;

[0055] Figure 15 illustrates a typical flight duty cycle for a delivery drone in operation;

[0056] Figure 16 illustrates how multiple battery cells / chemistries could operate in the context of the duty cycle illustrated in Figure 15,

[0057] Figure 17 is an overview of the control method for a drive unit and flight control system; Figure 18 is an exemplary Electronic Speed Controller for the drive unit of Figure 17, and

[0058] Figure 19 is an illustration of a delivery drone in flight.

[0059] It should be noted that the figures are not necessarily drawn to scale and that elements of similar structures or functions are generally represented by like reference numerals for illustrative purposes throughout the figures. It also should be noted that the figures are only intended to facilitate the description of the various embodiments described herein. The figures do not necessarily describe every aspect of the teachings disclosed herein and do not limit the scope of the claims.

[0060] Detailed Description

[0061] Before the present subject matter is described in detail, it is to be understood that this disclosure is not limited to the particular embodiments described, as such may, of course, vary. It is also to be understood that the terminology used herein is for the purpose of describing particular embodiments only, and is not intended to be limiting, since the scope of the present disclosure will be limited only by the appended claims.

[0062] It will be appreciated that the term rotor has two different applications in the context of an electrically powered aircraft. The first is that the aircraft is a rotorcraft with a rotor in the form of rotary wings or blades (rotor blades). The second is in the context of an electrical motor employed in the aircraft which may have an electrical rotor. To avoid confusion, the description below employs the term aero-rotor for the first meaning (i.e. rotor blades) with the term rotor adopted for an electrical rotor in an electric motor.

[0063] Drone delivery by necessity requires flying over populated areas which introduces a need for reliability, safety, and other characteristics such as low noise.

[0064] At the same time, commercial constraints require that drones can be loaded and unloaded quickly and fly at speed so that the time per delivery is minimised.

[0065] Existing designs for drone delivery are varied with several different approaches. Some designs employ multiple aero-rotors. For safety reasons, the number of aerorotors is generally selected to provide for redundancy.

[0066] As a result, typical configurations include a quad-coptor design with dual contrarotating aero-rotors or an octo-coptor.

[0067] However, these designs have demonstrated issues with respect to stability, reliability, and recovery in the event of a motor failure.

[0068] The present application provides an unmanned aerial vehicle (drone) which maintains the compact design, and superior drag coefficient of a quadcopter while increasing redundancy, control, and efficiency over a standard quadcopter design.

[0069] Representative examples of the embodiments described herein, which examples utilize many of these additional features and teachings both separately and in combination, will now be described in further detail with reference to the attached drawings. This detailed description is merely intended to teach a person of skill in the art further details for practicing preferred aspects of the present teachings and is not intended to limit the scope of the invention. Therefore, combinations of features and steps disclosed in the following detail description may not be necessary to practice the invention in the broadest sense, and are instead taught merely to particularly describe representative examples of the present teachings.

[0070] Moreover, the various features of the representative examples and the dependent claims may be combined in ways that are not specifically and explicitly enumerated in order to provide additional useful embodiments of the present teachings. In addition, it is expressly noted that all features disclosed in the description and / or the claims are intended to be disclosed separately and independently from each other for the purpose of original disclosure, as well as for the purpose of restricting the claimed subject matter independent of the compositions of the features in the embodiments and / or the claims. It is also expressly noted that all value ranges or indications of groups of entities disclose every possible intermediate value or intermediate entity for the purpose of original disclosure, as well as for the purpose of restricting the claimed subject matter.

[0071] The application will now be explained with reference to an exemplary delivery drone as shown in Figure 1 .

[0072] The exemplary drone 100 comprises an aircraft body 102. The drone may be defined with respect to its normal direction of forward flight to define a front side 104, rear side 106, left side 108, and a right side 110. The aircraft body has a top surface and an opposing bottom surface, which are connected by the sides.

[0073] The aircraft body may be shaped to reduce aerodynamic drag.

[0074] A landing gear may extend downwards from the bottom surface or the bottom surface may define a landing surface for the aircraft.

[0075] A plurality of arms 112, 114, 116, 118 extend from the aircraft body. Each arm supports at least one associated aero-rotor R1 , R2, R3, R4. Each aero-rotor is driven by an associated drive unit.

[0076] The aircraft body suitably houses a payload delivery system, a flight control system, at least energy supply system and any associated sensors. These sensors for example may include accelerometers, gyroscopes, magnetometers, and barometers.

[0077] The orientation, altitude and velocity of the drone are all controllable by individually adjusting the rotation speed of each of the four aero-rotors.

[0078] The payload delivery system is designed to allow for a payload to be loaded into the drone at a delivery base for delivery to a customer. Upon arrival at the customer’s location, the payload delivery system is configured in response to a command to release the payload to the customer.

[0079] As described below, the payload delivery system suitably allows the payload to be lowered to the ground by a tether connected to an electric motor (winch). Depending on the mode of operation, the tether may be released or the payload may be released by a mechanism and the tether then retracted back into the aircraft. The release mechanism may be passive where no actuator is involved (e.g. a simple hook). It may also be an active mechanism where an actuator causes the payload to be released. It will be appreciated that the person to whom the product is being delivered may also release the product.

[0080] As described below, the drone is configured to interact with an external control system. The system comprises the delivery ordering system, pilot / mission controller interface, the automatic flight planning system, the deconflicting system, sense and avoid algorithms, and fleet management functions.

[0081] In contrast to prior approaches, the present application does not seek to employ a redundant aero-rotor approach to provide for redundancy as interalia the redundant aero-rotor approach suffers from a loss of net aero-rotor aerodynamic efficiency from the co-axial arrangement due to the interaction of both aero-rotors wakes.

[0082] Instead, the present application provides a redundant motor design in which at least two electric motors are arranged to co-operate to drive a common aero-rotor. It will be appreciated that the arrangement is not restricted to a two motor configuration and that other configurations are possible. As an example, three motors could be provided sharing a common aero-rotor shaft. In this arrangement, two of the three motors could be configured to together provide the maximum operational torque required in the event of failure of the third motor.

[0083] It will be appreciated that for a drone to operate without flight control surfaces or auxiliary control aero-rotors, a minimum of three aero-rotors is required. It will be appreciated however that in the case of three rotors yaw control is not possible. Accordingly, the present application is explained in the context of a drone with four aero-rotors.

[0084] In the exemplary aircraft shown four arms extend from the aircraft body.

[0085] For a compact aircraft profile and other reasons, it is preferable to have only one aero-rotor provided on each arm.

[0086] The aero-rotor may have one or more aero-rotor blades arranged about a common axis of rotation. The rotation of each aero-rotor defines a plane of rotation. Suitably, the planes of rotation of the respective aero-rotors are parallel to one and other. Although, the planes may also be tilted with respect to one and other. Suitably, the planes of rotation of the aero-rotors on the front arms of the aircraft are in a common plane and the planes of rotation of the aero-rotors on the rear arms of the aircraft are in a common plane. The planes of rotation of the front arms and rear arms may be parallel with one and other but offset.

[0087] Preferably, the drone is configured to operate without the need for separate control surfaces. Instead, the flight of the drone is controlled by altering the rotation speed of the individual aero-rotors. It will be appreciated that depending on the number of aerorotors present, the precise manner of control will vary.

[0088] An exemplary arrangement of four aero-rotors is generally preferred as the method of flight control is relatively straightforward whilst at the same time maintaining a compact footprint for the drone.

[0089] With such an arrangement of four aero-rotors, two of the aero-rotors are configured to spin in a clock-wise direction (CW) and two of the aero-rotors are configured to spin in a counter-clock wise direction (CCW). One clock-wise rotating aero-rotor is positioned on a front arm of the drone with a counter clock-wise rotating aero-rotor positioned on the other front arm of the drone. The clockwise rotating aero-rotors are positioned on diagonally opposed arms to one and other and likewise the counter clock wise rotating aero-rotors are positioned on diagonally opposed arms to one and other.

[0090] A method of flight control 150 for such a four aero-rotor drone is illustrated in Figure 2, with the different control actions of altitude, yaw, pitch and roll shown. This method of control will be familiar to those skilled in the art. The control actions are effected by altering the speed of rotation of individual aero-rotors relative to the other aerorotors. In the illustration, the arrows upwards represent an increase in lift (resulting from an increase in aero-rotor speed) relative to another aero-rotor shown with a downward arrow having a corresponding decrease in lift.

[0091] Accordingly, for an increase in altitude, it may be seen that an increase in speed for each of the four aero-rotors results in an upwards motion of the drone (indicated by the arrow from the drone body). Equally, a decrease in altitude may be effected by a reduction in speed in each of the four aero-rotors.

[0092] Yaw (rotation about a vertical axis through the drone body) control may be effected by reducing the speed of the aero-rotors rotating in one direction and increasing the speed of the aero-rotors rotating in the other direction, which in turn will result in a rotation of the drone body about a vertical axis. Whilst, a simple reduction in speed in the aero-rotors rotating in direction will result in a yaw, the corresponding increase in speed of the aero-rotors rotating in the opposite direction is required to ensure that the altitude of the drone remains constant. Pitch and roll may be effected by altering the relative speeds of diagonally opposed aero-rotors. With pitch being about one diagonal and roll the other diagonal.

[0093] A system diagram for an exemplary drone is shown in Figure 3. The system 200 is arranged in a number of systems / elements with each system / element being directed to different functions of aircraft operation. It will be appreciated that these functions may be distributed and handled by more or less units than shown. One set of these functions may be grouped together as flight electronics 220. The flight electronics comprise a flight control system 206, a vehicle management system 208, and a communications system 210.

[0094] A Payload Retention and Delivery System 202 is responsible for receiving and delivering a payload. The vehicle management system provides overall management of the drone. The vehicle management system interfaces with the communications system to send and receive data to an off-board management system 212.

[0095] In a scenario with multiple drones operating from multiple launch sites, the off-board management system 212 may comprise one or more command units 214. In this scenario, each launch site may have a command unit which has a communications system for communicating with the individual drones. A management system 216, which may for example, be cloud based provides for the receipt of customer orders, processing those orders and scheduling drone deliveries accordingly.

[0096] The management system may also provide an interface to allow a drone pilot to create a flight plan for an individual delivery. This may be done using digital maps and other tools so as to ensure that a landing zone selected by a customer for their delivery is suitable and that the route for the flight is suitable and avoids any obstacles. Once a flight plan has been approved and a product is ready for dispatch to the customer, the management system may instruct the command unit with the flight plan. The command unit may in turn communicate this flight plan to a waiting drone. Suitably, the flight plan details the route the drone is to follow. The route may include altitude information. A flight dispatcher may then load the drone with the product for delivery and dispatch the drone.

[0097] The communications system may use any suitable RF communication technology to communicate with the command unit 214 of the off-board management system 216. Suitably, the communications are encrypted to prevent unauthorized parties from interfering with the operation of the drone.

[0098] A flight control system is provided to operate the drive units to maneuver the unmanned aircraft to a desired delivery location. A communications system provides at least one data connection to a remote (off-board) Management System. This Management system in turn provides for the receipt of delivery or collection instructions from customers. It will be appreciated that the management system may be responsible for the operation of multiple drones. The management system through a command unit may generate a flight plan for a particular flight operation to deliver or collect a product from a desired delivery / collection location. This flight plan is then communicated to the communications system of the drone which in turn delivers it to the vehicle management system and the flight control system.

[0099] A positioning system is provided for accurately determining the position of the drone in flight. The positioning system may comprise one or more receivers for a satellite positioning system such as GPS, Galileo, Glonass, BeiDou or QZSS.

[0100] The positioning system may be provided as part of the flight control system, the vehicle management system, or communications system.

[0101] The flight control system uses position measurements from the positioning system to effect control of the drive units to ensure that the drone follows the flight plan to the desired delivery location. It will be appreciated that the drone may also generate its own flight plan based on a point of departure and a point of collection / delivery. Once the drone has collected / delivered the product, the drone may follow the flight plan in reverse. Alternatively, it may be directed on a new flight plan which may be to a different destination. The flight control system may also be configured to land the drone or fly it back to base in the event of a failure being detected during flight.

[0102] The vehicle management or flight control systems may comprise a number of sensors for detecting dangerous situations, e.g. damage or failure of a motor and in response to detecting such a situation, be configured to return to the drone base or to effect an immediate landing. The choice of which may be determined based on the severity of the detected situation. The flight control system may also be configured to switch off all drive units and deploy a parachute in the event of an emergency arising during flight.

[0103] The powertrain comprises the plurality of drive units and one or more associated energy supply / storage systems.

[0104] The Payload Retention and Delivery (PRAD) System, 202, as shown in Figure 4, comprises a PRAD controller which is responsive to control commands from the Vehicle Management System to collect and deliver a payload. The primary actuator of the PRAD system is a winch motor 242 which is controlled by the PRAD controller through an electronic speed controller (ESC) 244.

[0105] In operation, the payload is connected by a tether which is wound onto a drum. The drum in turn may be wound / unwound by the winch so that the payload may be lowered or raised as required from the drone. In use, the drum with the tether may loaded at the same time as the payload. The payload may be retained in flight in a cargo bay.

[0106] The cargo bay may have doors which are latched retaining the payload. At a desired delivery location, the PRAD controller may operate a latch actuator 246 to cause the cargo bay door(s) to open. At which time the PRAD controller can operate the winch to lower the payload to the ground.

[0107] A line cutter 248 may be provided to cut the tether. The line cutter may be activated to cut the tether in the event of an emergency or once a payload has been delivered to the ground.

[0108] A load cell 250 may be connected, through an analog to digital converter 252, to the PRAD controller. The load cell suitably indicates the load on the tether / winch. In the event of an overload situation, e.g. the tether having become entangled, the PRAD controller may actuate the line cutter so that the safe operation of the drone is ensured.

[0109] Equally, the load cell may be used to identify when the product has landed on the landing site, by virtue of the load effectively reducing to zero when the product rests on the landing surface.

[0110] A tensioner actuator may be provided to maintain the tension in the tether relatively constant. The tensioner actuator is useful to maintain tension on the tether after product delivery when the tether is being re-wound on the drum of the winch. An encoder 252 may be provided on the winch so that the PRAD controller is aware of how much of the tether has been unwound, i.e. how far the payload has been lowered from the drone. This allows the PRAD to operate the line cutter once the payload has arrived on the ground or to determine how much tether remains to be rewound after product delivery.

[0111] A focus of the present application is that of the drive units for the aero-rotors. Product delivery demands that drones operate in all sorts of weather.

[0112] The bandwidth (response speed) of the propulsion unit (aero-rotors) is the dominant factor in a drone’s agility, manageability, and disturbance rejection. The faster the response is, the better a drone will perform in adverse weather conditions. Field trials of dual aero-rotor contra-rotation drones by the applicant have identified significant drawbacks in the operation of this configuration.

[0113] Arising from this, it was determined that a different approach was required.

[0114] This has resulted in the development of the drive units in the present application, in which, multiple motors are employed in each drive unit to drive a common aero-rotor.

[0115] As described below, a dual motor single aero-rotor unit, sized for one-motor- inoperative case (hence capable of ~double the required normal operational torque) is provided.

[0116] This arrangement has a superior performance in terms of bandwidth relative to two independent units, each driving a separate aero-rotor, since the bandwidth is mostly dominated by the maximal angular acceleration of the shaft: il = — , where il is

[0117] I XX angular acceleration, Q is the torque and Ixx is the sum of the inertia of the motor and the aero-rotor. As the rotor inertia is constant, doubling the torque (Q) will result in increasing the acceleration hence the bandwidth.

[0118] An additional advantage of a single aero-rotor per arm configuration is a significant reduction in the aero-rotor drag force, which is proportional to the aero-rotor side cross section.

[0119] At the same time, having a single aero-rotor vs. dual aero-rotor configuration means approximately half the aero-rotor drag. Thus resulting in a higher flight speed and improved range for a drone.

[0120] Another advantage of the single aero-rotor configuration is the ability to seamlessly recover from a one-motor-inoperative failure case.

[0121] In particular, in the case of a single aero-rotor-dual motor configuration, the failure identification and reconfiguration is taking place at the propulsion unit level. In contrast in other redundant configurations, where failure identification is presented at the air vehicle level.

[0122] As a result, the response time, relative to other configurations, is extremely short. When this short response time is combined with the RPM drop rate (due to the motor and aero-rotor inertia), the vehicle attitude and location will stay (almost) as before the failure occurrence. In contrast, in other configurations, there is a tendency to significantly drift from their planned flight path when encountering such a failure, increasing the risk of collision with obstacles, the ground or other drones.

[0123] A first exemplary arrangement of a drive unit 300 for an aero-rotor with a redundant motor design is shown in Figure 5.

[0124] In this exemplary arrangement each drive unit comprises two motors 302, 304. The two motors are arranged to be co-axial with one and other along a common rotational axis 308. The two motors share a common rotor shaft 306. The rotor shaft may be a single piece or may be comprised of separate rotor shafts connected by a connecting element C1.

[0125] The motors are configured to rotate in a common direction (either CW or CCW).

[0126] The electric motors may be brushed or brushless motors. Brushless motors are generally preferred as they are more efficient, more powerful for their weight, and have a longer operating life. The operation of electric motors will be familiar to those skilled in the art, with each motor having a rotor part R1 , R2 and a stator part S1 , S2 where the energizing of windings causes magnetic forces to be set-up between the rotor and stator which in turn causes the rotation of the rotor. Bearings B1 , B2, B3, B4 are provided to allow the relative rotation between the rotor and stator parts.

[0127] In the exemplary arrangement shown, when the motor windings provided as part of the stators S1 , S2 become energized, a temporary magnetic field is created that repels (and / or attracts) against the permanent magnets present inside the motor on the corresponding rotor R1 ,R2.

[0128] The motors’ rotors (R1 , R2) are connected by a connecting rotor shaft (C1 ). The rotor blades are attached to one end of the rotor. For ease of illustration, the propeller mount on the top rotor, R1 , is not shown. The rotor blades may be permanently fixed to the rotor shaft or may provide for a releasable fixing. This may for example be a quick release mechanism as would be familiar to those in the field of drones.

[0129] By virtue of the connecting rotor shaft, either or both of the motors (S1 , S2) may be operated to rotate the rotor blade.

[0130] In the exemplary arrangement, the stators S1 , S2 of the first and second motors 302, 304 are connected by a motor housing (H1 ). The rotor shafts are supported by the bearings (B1 , B2, B3, B4) spaced along the rotational axis.

[0131] Each drive unit is configured to provide a maximum operational torque, MOT, to the rotor. To provide redundancy, each of the first and second motors are suitably configured to individually provide a torque sufficiently close to the maximum operational torque to maintain effective flight. Thus in the event of the failure of one motor, the remaining motor has sufficient reserve available to assume the load for both motors.

[0132] In a simple example with two motors, each motor can provide the MOT. In practise, to save weight, each motor may provide a factor a of the MOT. The lower limit of a in a two motor configuration would be .5, i.e. both motors together are required to provide the maximum operational torque. The upper limit being 1 or more, i.e. a single drive unit can provide the MOT. Thus providing a range for a given by ,5< a<1 . Suitably, a is greater than .7. More suitably, a is greater than .8.

[0133] It will be appreciated that the approach is not restricted to two motors per drive unit (aero-rotor) and that it is possible to employ three or more motors. With three or more motors, the individual motors need not be required to provide the maximum operational torque since in the event of the failure of one motor, the remaining motors would share the load. Thus in the case of N motors sharing a maximum operational torque for a drive unit that the individual maximum operation torque for each motor would be: a M0T / (N-1 ), where N is the number of electric motors in a drive unit and a represents a design factor corresponding to what fraction of the MOT is required to be produced in the event of the failure of one of the motors.

[0134] The motors need not necessarily be arranged along a common rotational axis but may instead to be connected together by a gearbox or other arrangement. An example of such an arrangement 350 is shown in Figure 6, in which two motors 352, 354 are each positioned to drive a flywheel 356. The flywheel in turn is connected to the propeller of the aero-rotor. The drive between the motors and flywheel may, for example, be by friction or toothed gears about the motor axis and flywheel axis.

[0135] A further approach is to integrate two or more motors into a single structure. In such an arrangement, a motor is provided with two or more separate sets of stator windings. In this arrangement, the motor stators may be wound together as different segments on a segmented stator. A first set of windings is driven by one ESC with the second set of windings driven by a second ESC. In this arrangement, each set of windings is configured to generate an electromagnetic force to cause the rotor of the motor to rotate relative to the stator of the motor. In this arrangement, the windings may be regarded as two separate motors as each set of windings may be independently controlled to cause rotation. Equally, they may be concurrently energised to also cause rotation.

[0136] Whilst reliability is one design factor, serviceability is another as any delays putting a drone back into service will have an effect on product deliveries. Accordingly, it is desirable that drones are designed with elements that may be replaced quickly. Addressing this, Figure 7 depicts an exemplary implementation 400 of a dual redundant motor design incorporated into Releasable Drive unit which allows for quick removal and replacement.

[0137] In this exemplary drive unit 400, as before there are two motors M1 , M2 for example of the type generally described above.

[0138] The two motors have a common drive shaft to which the aero-rotor blade (R1 ) is mounted. The two motors are in turn mounted onto a housing H1. The housing in turn contains an electronic speed controller ESC1 , ESC2 for each of the individual motors M1 , M2. The two electronic speed controllers may be provided as a single dual electronic speed controller. The housing in turn is configured to connect to an arm A1 of the drone. The method of connection is one which provides mechanical strength and electrical connections.

[0139] For example, the connection may be provided by means of two quick connectdisconnect connectors (C1 , C2), that allow signals and power to pass from the housing to the arm A1 of the aircraft. Alternatively the housing may define an arm of the aircraft and the connection is made directly to the aircraft body.

[0140] It will be appreciated that having the ESCs co-located with the motors in the releasable drive unit means that it is not necessary to perform fault analysis on the drone to determine whether a failure has arisen in a motor or the associated ESC. Instead, the entire unit can simply be replaced and the faulty unit analysed later to identify and potentially fix the fault. This is a significant advantage where continuous flight operations are desired such as in a product delivery environment.

[0141] An exemplary electric propulsion architecture 450 for a dual motor quadcopter for drone delivery service with a single power supply system is shown in Figure 8.

[0142] In this arrangement, each drive unit (motor assembly) comprises a pair of electric motors (M11-M12, M21-M22, M31 -M32, M41-M42) and an aero-rotor (R1 , R2, R3, R4).

[0143] The aero-rotor is mechanically coupled to both motors so that during operation either or both motors of the pair is operable to drive the aero-rotor blades. The electric motors may be direct drive as per figure 5, i.e. not requiring a gear box to operate the aero-rotor at a desired speed. Alternatively, the motors may be connected by a gear box or flywheel per figure 6.

[0144] Each motor has a corresponding electronic speed controller, ESC, (ESC11 , ESC12, ESC 21 ,ESC22, ESC 31 , ESC32, ESC 41 , ESC42) which provides electrical signals to control the speed of the motor in response to signals from the flight control system. It will be appreciated that typically an electronic speed controller converts a DC voltage to an AC voltage which is used to energise the windings of an electric motor. An electronic speed controller may control the current, torque, and / or speed of an electric motor.

[0145] This exemplary architecture comprises one power supply system that provide power to each of the 8 ESC’s. The power supply system in turn comprises an energy storage system, ESS. The power supply system suitably comprises conductors for providing electrical power to the ESCs. These conductors may be arranged as a combination of one or more bus bars, cabling, and connectors. The ESS may include protection features, such as fuses, to ensure the ESS is not accidentally short circuited.

[0146] The ESS comprises one or more energy storage elements. The storage elements, may, for example, be batteries. The batteries may, for example, be lithium polymer batteries, lithium-ion batteries or nickel-cadmium batteries.

[0147] Alternatively, the energy storage elements may comprise fuel cells. Suitably the ESS is arranged in an architecture where the failure of a single energy storage element does not result in the failure of the entire ESS.

[0148] Accordingly, an exemplary architecture employed for the ESS is a multi-string solution. In such a multi-string arrangement 500, a plurality of batteries are arranged in a parallel - series arrangement such as shown in Figure 9. In this arrangement, the ESS is arranged as several parallel strings of serially connected energy storage elements. In the event of the failure of one of the strings, e.g. by the failure of one of the energy storage elements in the string, the remaining strings can deliver the required load.

[0149] An alternative arrangement 550, shown in Figure 10, parallels a plurality of energy storage elements before putting them in series. Each of the energy storage elements has a circuit protection feature, such as a fuse. The circuit protection circuit acts to disconnect an individual energy storage element in the event of a failure preventing any thermal event.

[0150] The multi-string solution is capable of managing a full string / cell failure. The string will be disconnected and there is enough energy in the remaining strings to return home in the worst case conditions. The enclosure is designed to prevent any thermal events that have occurred from spreading to any other strings for at least the time it takes for the drone to return home.

[0151] Suitably, the ESS is mechanically arranged such that baffles, walls or other structures are in place to ensure that a thermal failure (overheating or potentially fire) in one energy storage element or string does not affect or spread to the other energy storage elements or strings.

[0152] The arrangement of strings and cells ensures that there is redundancy in the event of a failure of a string or cell as is illustrated in Figures 11 and 12. In the exemplary operating case 600, as shown in Figure 11 , each drive unit is operating at 100% of maximum, with each individual motor in each drive unit contributing an equal share of 50% of the drive to the aero-rotors.

[0153] In the event of a motor failure scenario 650 (in motor M12 in the drive unit for the first aero-rotor R1 ), as illustrated in Figure 12, the ESC (ESCH ) for the remaining motor M1 1 increases the output to its maximum, in this case 80% (i.e. the value of a is .8). To compensate for this reduction in output on the first aero-rotor, the remaining drive units through their ESC’s reduce their maximum output to 80% (40% from each motor) to compensate. It will be appreciated that as only the maximum power that can be delivered is reduced, the control functions for the drone are unaffected.

[0154] Under this reduced maximum operating power, the drone is suitably configured to return to base and other than potentially flying at a reduced speed, the operation of the drone will be unaffected.

[0155] The drone may also be configured to operate a redundant power supply system, in which there are two distinct energy storage systems.

[0156] Such an architecture is similar to the previously discussed architecture but comprises of two energy storage systems that provide power to the 8 ESC’s. More specifically, as shown in Figure 13. In this architecture 700, as before, there are four aero-rotors (R1 ;R2;R3;R4) with each aero-rotor driven by a drive unit having two motors (M11.M12; M21 , M22;M31 ,M32; M41 , M42).

[0157] Two Energy Storage Systems (ESS1 , ESS2) are provided. The first energy storage system (ESS1 ) provides power to one motor in each of the drive units, in the example there are M12, M22, M32, M42. The second energy storage system (ESS2) provides power to the remaining four motors (M11 , M21 , M31 , M41 ) in each of the drive units.

[0158] This approach provides for a number of advantages.

[0159] A first advantage of this approach is illustrated in the scenario 750 of Figure 14, in which one of the ESS’s (ESS2) has failed. In the earlier arrangements, the failure of the single ESS would have rendered the drone non flyable necessitating for example the deployment of a parachute. However, in the arrangement of Figure 14, the remaining ESS provides power to one motor of each drive unit so that the drone may remain flying. As before, each motor in a drive unit is selected to be able to provide more than 50% of the required maximum so that in the event of failure, the drone is operable at over 50% of maximum output. In the exemplary arrangement shown, each motor is capable of operating at 80%. As in the failure case of Figure 12, suitably the drone would return to base. It will be appreciated that the dual ESS arrangement can similarly account for a single motor failure as the single ESS arrangement.

[0160] A further advantage of the dual ESS architecture is that a higher levels of redundancy is provided for other elements apart from energy storage, e.g. with respect to connectors and bus bars connecting each ESS to the ESC’s.

[0161] A further advantage of having two separate ESS’s is that it facilitates the use of different cell types / chemistries.

[0162] More particularly, one ESS may be provided with one cell type / chemistry with the remaining ESS provided with a different cell type / chemistry.

[0163] This approach allows for more efficient operation during flight.

[0164] This may be explained by reference to Figure 15 which illustrates a typical duty cycle 800 for current delivered to a drive unit during a flight of just under 450 seconds.

[0165] As may be seen, there is a higher current during take-off and flight to the point of delivery of the product. After which with a reduced weight, the drone requires less lift from the aero-rotors to fly.

[0166] It will be appreciated that batteries are generally optimised for a particular operating scenario. However, as may be seen from the typical duty cycle of Figure 16, the conditions can vary significantly.

[0167] To account for this, a first ESS could be selected to comprise a lower C-rate (a measure of the rate at which a battery is discharged relative to its capacity), higher energy cell that is efficient at providing low current. Thus, in the proposed power sharing approach 850, illustrated by the dotted line in Figure 16, the first ESS is configured to provide current at 40A or lower over the whole duty cycle. In contrast, the second ESS could use a higher C-rate, but lower energy density cell that could power the high transients seen during the duty cycle (for example anything above 40A).

[0168] It will be understood, that each of the ESS could operate in the territory of the other, for example in the event of a failure of an ESS, for the limited time necessary to return to base.

[0169] Similarly, the characteristics of the motor may be selected to match the chemistry / configuration of the associated ESS. As a simple example, one ESS may be configured to deliver a low voltage - high current and the other ESS may be configured to deliver a high voltage - low current. It will be appreciated that in such a scenario, one motor of each pair will be selected to a low voltage motor and the other may be selected to a high voltage motor. It will be appreciated that low and high voltage are with respect to one and other. Thus for example, a low voltage might be considered to be 12V and a high voltage 24 or 48V.

[0170] Typically, power is shared 50:50 between both motors to drive each aero-rotor. In the event of a failure, the redundant motor (and associated power architecture) can allow the drone to return home safely in a degraded state. Each motor is capable of providing more than 50% of the thrust.

[0171] More advanced power sharing is also possible. Depending on the load point or thermal conditions it may be preferable to share the power in a different ratio than 50:50 between the motors in each drive unit.

[0172] Such an approach may improve efficiency. It may also improve the thermal performance of the drive units. In this respect, it will be appreciated that the operating life of components is affected by their temperature, as set forth by the Arrhenius relationship. Accordingly, adjusting the ratio between motors may be useful to ensure that a circuit (e.g. ESC) or component (motor) does not overheat. Thermal sensors may be deployed to monitor the temperatures of the motors and / or ESC.

[0173] An exemplary arrangement 900 for a drive unit 900 is shown in Figure 17 to account for the modes of operation described above.

[0174] In this arrangement, the flight control system, sometimes referred to as a FCU, communicates both speed demand and power ration, PR, to both ESC’s of each drive unit.

[0175] To operate effectively, the two ESC’s may be configured in a master and slave configuration.

[0176] In this configuration, one of the ESCs is selected to initially be the master. This may be by default, i.e. ESC1 is selected as master during start-up. Alternatively, there may be an arbitration process to select the master. Such processes would be familiar to those skilled in the art. The master arbitrates between both ESC and ensure they don’t work against each other when controlling speed. If the master ESC fails the other ESC is configured to automatically assume the role of master.

[0177] As both motors can be active at the same time this approach doesn’t suffer from complex control of modes.

[0178] In contrast to previous approaches, the use of dual ESC-dual motors have significantly improved recovery time, with a demonstrated recovery in less than ~200 msec meaning that any altitude loss or position drift is typically less than 0.5 m. It will be appreciated that these figures will depend on a number of factors.

[0179] An exemplary electronic speed controller, ESC, 920 for use in the arrangement of Figure 17 is illustrated in Figure 18 and suitably comprises a micro-controller, gate driver(s) and semiconductor switch(s).

[0180] The micro-controller receives control signals from the flight controller. In this configuration, these signals are a speed (power) control value and a power ratio value. The first indicates how much effort is required to be provided overall by the two motors with the second indicating the share between the Master and the Slave.

[0181] The micro-controller interacts with the other micro-controller in the other ESC of the same drive unit to share the power as dictated by the power ratio value. In the event of a non-response from the other ESC, the micro-controller can assume the full load. The microcontroller may also provide signals back to the flight control system. Such signals may identify a fault and / or operating conditions (e.g. current, temperature etc).

[0182] The microcontroller generates one or more switching signals which are passed to one or more gate drivers (depending on the switching configuration). The one or more gate drivers are in turn employed to control semiconductor switches which connect power to drive the motor. The semiconductor switches may be MOSFETs or other suitable transistors. Typically, the MOSFETs are switched to convert a DC voltage into an AC (or quasi AC) voltage which is applied to the winding of a motor.

[0183] In another configuration, a single microcontroller may be shared by both ESCs with the single microcontroller providing control signals to the gate drivers of each ESC which in turn provide switching signals to the semiconductor devices controlling each motor. It will be appreciated that this saves on a microcontroller but forgoes redundancy in the event of the failure of the microcontroller. This arrangement may also be considered as a single ESC with dual motor outputs.

[0184] The flight control system and vehicle management system employed in the present system may be of the type generally employed within drones with appropriate modifications made to account for the control of the unique aspects presented herein including for example the use of dual motors in the drive units and the PRAD.

[0185] As an example, the flight control system may comprise a single flight controller. Alternatively, it may comprise multiple flight controllers. With multiple flight controllers, a voting or other mechanism may be employed to determine which flight controller acts to control the operation of the drone. An exemplary illustration of a drone 950 performing a product 958 delivery to a delivery location 960 is shown in Figure 19. The exemplary drone has 4 aero-rotors 954, each in turn driven by a drive unit 956. The drive units are mounted on arms 962 which extend from the aircraft body 952. The product is suspended by a tether below the aircraft body.

[0186] While the present application has been described in the context of drone delivery, it will be appreciated that it may have application in other aircraft not employed for drone delivery but where safety is critical, for example, an aircraft being used over populated areas. The use may be for photography or sensing. It will be appreciated that in such applications, the product delivery module may be substituted with a module with the requisite cameras or sensors for the purpose.

[0187] The example embodiments provided herein, however, are merely intended as illustrative examples and not to be limiting in any way.

[0188] All features, elements, components, functions, and steps described with respect to any embodiment provided herein are intended to be freely combinable and substitutable with those from any other embodiment. If a certain feature, element, component, function, or step is described with respect to only one embodiment, then it should be understood that that feature, element, component, function, or step can be used with every other embodiment described herein unless explicitly stated otherwise. This paragraph therefore serves as antecedent basis and written support for the introduction of claims, at any time, that combine features, elements, components, functions, and steps from different embodiments, or that substitute features, elements, components, functions, and steps from one embodiment with those of another, even if the following description does not explicitly state, in a particular instance, that such combinations or substitutions are possible. Express recitation of every possible combination and substitution is overly burdensome, especially given that the permissibility of each and every such combination and substitution will be readily recognized by those of ordinary skill in the art upon reading this description.

[0189] In many instances, entities are described herein as being coupled to other entities. It should be understood that the terms “coupled” and “connected” (or any of their forms) are used interchangeably herein and, in both cases, are generic to the direct coupling of two entities (without any non-negligible (e.g., parasitic) intervening entities) and the indirect coupling of two entities (with one or more non-negligible intervening entities). Where entities are shown as being directly coupled together, or described as coupled together without description of any intervening entity, it should be understood that those entities can be indirectly coupled together as well unless the context clearly dictates otherwise.

[0190] While the embodiments are susceptible to various modifications and alternative forms, specific examples thereof have been shown in the drawings and are herein described in detail. It should be understood, however, that these embodiments are not to be limited to the particular form disclosed, but to the contrary, these embodiments are to cover all modifications, equivalents, and alternatives falling within the spirit of the disclosure. Furthermore, any features, functions, steps, or elements of the embodiments may be recited in or added to the claims, as well as negative limitations that define the inventive scope of the claims by features, functions, steps, or elements that are not within that scope.

Claims

Claims1. An unmanned aerial vehicle for product delivery comprising: an aircraft body; a plurality of aero-rotors provided about the aircraft body, each aerorotor being arranged to rotate to provide lift for the aircraft; a payload delivery system for delivering a payload at a desired delivery location; a plurality of drive units, each drive unit being configured to rotate an associated aero-rotor, with each drive unit comprising a plurality of electric motors, wherein each individual electric motor of the plurality of electric motors alone or in combination with the other electric motor(s) of the plurality is operable to rotate the associated aero-rotor; at least one energy supply system providing power to the drive units; and a flight control system for controlling the aircraft, the flight control system being operable to operate the drive units to manoeuvre the unmanned aircraft to a desired delivery location.

2. The unmanned aerial vehicle of claim 1 , wherein each drive unit is designed to provide a maximum operational torque, MOT, and where each motor of the drive unit is configured to provide a maximum operational torque of a M0T / (N-1), where N is the number of electric motors in a drive unit and a is a factor determine a reduced maximum operational torque in the event of a failure.

3. The unmanned aerial vehicle of claim 2, wherein each motor has a motor rotor rotating about a rotor shaft.

4. The unmanned aerial vehicle of claim 3, wherein the rotor shafts of each motor are connected together.

5. The unmanned aerial vehicle of any one of claims 3 to 4, wherein the rotor shafts are co-axial.

6. The unmanned aerial vehicle of claim 3, wherein the motors share a common rotor shaft.

7. The unmanned aerial vehicle of claim 3, wherein the motors are cohoused and provided as a single unit, each motor comprising a separately excitable motor stator with the motors sharing a common motor rotor.

8. The unmanned aerial vehicle of claim 7, wherein the motor stators are wound together as different segments on a segmented stator.

9. The unmanned aerial vehicle of any preceding claim, wherein each drive unit is directly connected to an associated aero-rotor without the use of a gearbox.

10. The unmanned aerial vehicle of claim 3, wherein the rotor shafts of each motor are connected through a fly or gear wheel to the aero-rotor.

11. The unmanned aerial vehicle of any preceding claim, wherein each drive unit has a first motor and a second motor and where, the first and second motors in combination are configured to provide a maximum operational torque, MOT, to drive the aero-rotor and wherein each of the first and second motors is individually capable of delivering a proportion of the maximum operational torque, the proportion being greater than 70%.

12. The unmanned aerial vehicle of any preceding claim, further comprising an electronic speed controller associated with each motor of each drive unit, each electronic speed controller being configured in response to one or more signals from the flight control system to control the power delivered from the at least one electrical power source to the associated motor.

13. The unmanned aerial vehicle of any one of claims 1 to 12, further comprising an electronic speed controller for each drive unit, the electronic speed controller being associated with all of the motors in a drive unit, the electronic speed controller being configured in response to one or more signals from the flight control system to control the power delivered from the at least one electrical power source to the motors of the drive unit.

14. The unmanned aerial vehicle of claim 12 or 13, wherein each drive unit comprises a housing.

15. The unmanned aerial vehicle of claim 14, wherein the one or more electronic speed controllers for a drive unit are integrated with the motors in the drive unit.

16. The unmanned aerial vehicle of claim 15, wherein the housing is connected by a releasable connection to the aircraft.

17. The unmanned aerial vehicle of claim 16, wherein the releasable connection is a quick connect-disconnect connection.

18. The unmanned aerial vehicle of any preceding claim, wherein there are a plurality of energy supply systems, each of the plurality of energy supply systems uniquely suppling energy to a motor in each of the driveunits so as to ensure that in the event of a failure of one energy supply system, that only those motors connected to that energy supply system will fail and the remaining motors can continue to operate being supplied by the remaining energy supply system.

19. The unmanned aerial vehicle of any one of claims 1 to 18, further comprising an energy supply system for supplying electrical energy to each drive unit.

20. The unmanned aerial vehicle of claim 19, wherein the energy supply system comprises a plurality of cells arranged as multiple series strings in parallel, where individual strings of the series can be electrically disconnected in the event of a failure.

21. The unmanned aerial vehicle of claim 19 wherein the energy supply system comprises a plurality of cells connected as multiple parallel strings arranged in series, where individual cells can be disconnected in the event of a failure.

22. The unmanned aerial vehicle of any one of claims 19 to 21 , wherein the energy supply system comprises cells of different battery chemistry arranged in a multi-battery chemistry pack providing a single output to each drive unit.

23. The unmanned aerial vehicle of claim 22, wherein a first battery chemistry is selected to provide a first cell with a first C-rate and having a first energy capacity and a second battery chemistry is selected to provide a second cell with a second C-rate and having a second energy capacity, wherein the first C-rate is lower than the second C-rate and the first energy capacity is higher than the second energy capacity.

24. The unmanned aerial vehicle of any one of claims 19 to 23, wherein the energy supply system comprises cells of different battery chemistry arranged in a multi-battery chemistry pack providing at least two outputs to each drive unit with each output powering a different motor of the drive unit.

25. The unmanned aerial vehicle of claim 24, wherein the motors of the drive unit have different operating characteristics with each motor being selected to match the battery chemistry providing the output to which the motor is connected.

26. The unmanned aerial vehicle of claim 11 , wherein one of the first and second motors is selected to be an active motor which in flight is powered to drive the aero-rotor with the remaining motor selected to bea passive motor which is unpowered.

27. The unmanned aerial vehicle of claim 26, wherein upon the failure of active motor, the passive motor is selected to assume the role of the active motor.

28. The unmanned aerial vehicle of claim 26, wherein the first and second motors, in flight, share the drive load substantially equally.

29. The unmanned aerial vehicle of claim 26, wherein the ratio of power delivered by each motor is varied dynamically during flight.

30. The unmanned aerial vehicle of claim 29, wherein the ratio of power is varied based on current flight conditions.

31. The unmanned aerial vehicle of claim 29, wherein the ratio of power is varied to account for thermal differences between motors or semiconductor devices controlling the motors.

32. The aircraft according to any preceding claim, wherein the payload delivery system for delivering a payload at a desired delivery location is configured to collect the payload from a sender’s location.

33. An unmanned aerial vehicle comprising: an aircraft body; a plurality of rotors provided about the aircraft body, each rotor being arranged to rotate to provide lift for the aircraft; a plurality of drive units, each drive unit being configured to rotate an associated aero-rotor, each drive unit comprising a motor having a rotor and at least two stators wound together on the same former; at least one energy supply system providing power to the drive units; and a flight control system for controlling the aircraft, the flight control system being operable to operate the drive units to manoeuvre the unmanned aircraft to a desired delivery location.

34. The unmanned aerial vehicle of claim 33, further comprising an electronic speed controller associated with each stator, each electronic speed controller being configured in response to a signal from the flight control system to control the power delivered from the at least one electrical power source to the associated stator.

35. The unmanned aerial vehicle of any one of claims 33 or 34, wherein there are a plurality of energy supply systems, each of the plurality of energy supply systems uniquely suppling energy to a stator in each of the drive units so as to ensure that in the event of a failure of oneenergy supply system, that only those stators connected to that energy supply system will fail and the remaining stators can continue to operate being supplied by the remaining energy supply system.

36. The unmanned aerial vehicle of any one of claims 33 to 35, wherein one of the first and second stators is selected to be an active stator which in flight is powered with the remaining stator selected to be a passive motor which is unpowered.

37. The unmanned aerial vehicle of claim 36, wherein upon the failure of active stator, the passive stator is selected to assume the role of the active stator.

38. The unmanned aerial vehicle of any one of claims 33 to 35, wherein the first and second stators, in flight, share the drive load substantially equally.

39. A drone, comprising: four aero-rotors; four drive units, each drive unit operable to rotate an associated aerorotor, each drive unit comprising two stators and at least one rotor; eight electronic speed controllers (ESC), each electronic speed controller converting a DC voltage to an AC voltage and providing the AC voltage to an associated stator.

40. A drone according to claim 39, wherein in any given drive unit, the stators may both be active or a single stator may be active.

41. A drone according to claim 40, wherein when both stators are active, the power split between the stators can be dynamically changed.

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