GB2569659A

GB2569659A - Airborne urban mobility vehicle with VTOL (Vertical Take-Off and Landing) capability

Info

Publication number: GB2569659A
Application number: GB1721837.1A
Authority: GB
Inventors: Didey Arnaud
Original assignee: Didey Consulting Ltd
Current assignee: Didey Consulting Ltd
Priority date: 2017-12-22
Filing date: 2017-12-22
Publication date: 2019-06-26
Also published as: GB201721837D0

Abstract

An airborne urban mobility vehicle 1 having vertical take-off and landing (VTOL) capability, comprises a fuselage 200 pivoted between lateral arms of a yoke 220, extending fore and aft and, at or towards the extremities of the arms; the respective fore portions are linked laterally together by an aerofoil 240; and the respective aft portions are linked laterally by an aerofoil 250; and at least one of the fore and aft aerofoils having a one or more propulsion units 260 mounted thereon. The aerofoils may have between one and four propulsion units mounted thereon and be equal between fore and aft, port and starboard. The aerofoils may comprise fixed wings in relation to the yoke, and the yoke moves with respect to the fuselage at the pivot (270, fig 2). The propulsion means may be controlled by a flight control unit which distributes electrical power to the propulsion units. The aircraft may be maneuvered in pitch, roll and yaw, and transition from vertical to horizontal flight by means of adjusting the relative propulsive force provided by the propulsion units. Rotation of the fuselage may be mediated to limit/enhance movement and may be acted on by a braking arrangement or actuator.

Description

(54) Title of the Invention: Airborne urban mobility vehicle with VTOL (Vertical Take-Off and Landing) capability Abstract Title: Airborne urban mobility vehicle with VTOL capability.

(57) An airborne urban mobility vehicle 1 having vertical take-off and landing (VTOL) capability, comprises a fuselage 200 pivoted between lateral arms of a yoke 220, extending fore and aft and, at or towards the extremities of the arms; the respective fore portions are linked laterally together by an aerofoil 240; and the respective aft portions are linked laterally by an aerofoil 250; and at least one of the fore and aft aerofoils having a one or more propulsion units 260 mounted thereon. The aerofoils may have between one and four propulsion units mounted thereon and be equal between fore and aft, port and starboard. The aerofoils may comprise fixed wings in relation to the yoke, and the yoke moves with respect to the fuselage at the pivot (270, fig 2). The propulsion means may be controlled by a flight control unit which distributes electrical power to the propulsion units. The aircraft may be maneuvered in pitch, roll and yaw, and transition from vertical to horizontal flight by means of adjusting the relative propulsive force provided by the propulsion units. Rotation of the fuselage may be mediated to limit/enhance movement and may be acted on by a braking arrangement or actuator.

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Application No. GB1721837.1

RTM

Date :25 June 2018

Intellectual Property Office

The following terms are registered trade marks and should be read as such wherever they occur in this document:
UBER	page 1
eHang	pages 1, 2
Volocopter	pages 1, 2
Airbus	pages 1, 2
ItalDesign	pages 1, 2
Boeing	page 2
Lillium	pages 2, 3

Intellectual Property Office is an operating name of the Patent Office www.gov.uk/ipo

Airborne Urban Mobility Vehicle with VTOL (Vertical Take-Off and Landing) capability

SCOPE

The present invention relates to an A-UMV (Airborne Urban Mobility Vehicle) with VTOL (Vertical Take-Off and Landing) capability. A-UMV is a development of the macro copter and drone concepts but scaled up so as to potentially carry a person or at least a more significant payload.

BACKGROUND

The term A-UMV is used to describe an airborne vehicle designed to enhance mobility in and around congested cities and metropolises, avoiding traffic jams (typically above-ground) or overcrowded public transportation systems (under and above-ground).

A-UMV available to travellers are currently essentially limited to helicopter typically used within large cities equipped for example with roof-tops landing pads. Helicopters are however very expensive to procure and maintain, noisy, bulky and on the whole reliant on non-renewable fuel, contributing to further emissions in already heavily polluted cities.

It is anticipated that the use of electric A-UMV will expand dramatically as congestion on roads in and around cities result in ever increasing commute/journey time. The business case for such future mode of transport is comprehensively discussed in UBER Elevate in what UBER refers to as on-demand aviation.

This future will only be made possible with the evolution of infrastructures, regulations and technologies. In particular, amongst key enablers of electric A-UMV are electric propulsion systems and in particular the batteries and charging technologies required to store electrical energy and top-up between flights. In addition, to ensure that A-UMV can be safely accessed to as many users as possible, sophisticated flight control algorithms will have to be developed to assist users/riders in piloting such aircraft before eventually enabling autonomous flight.

Whilst the above infrastructure is essential there is also requirement to provide an efficient AUMV concept which optimises the use of electric propulsion and can negotiate a crowded urban environment.

Known and Proposed A-UMV Systems

A number of novel electric A-UMV concepts are also currently under development, in an effort to negate the drawbacks of helicopters (complexity, noise, pollution, cost...). Two different types of A-UMV are being developed, Multi-copters and VTOL aircraft:

Multi-copters

Representative Multi-copter UMV’s, are those developed by the Chinese company eHang, the German company Volocopter or Airbus concept developed by ItalDesign.

The E Hang concept provides a central fuselage, on an effectively rectangular base, with arms protruding from the vertices of that base, the arms terminating in double sided propeller arrangements. The Volocoopter 2X concept provides a central helicopter like fuselage but in place of the helicopter rotor a network of beams supports a plurality, around 18 electrically driven propellers. The ItalDesign concept provides a central helicopter like fuselage above which for arms protrude, the arms terminating in ducted propellers.

These are essentially electric helicopters benefiting from the simplicity afforded by distributed fixed-pitch propellers. Such aircraft are ideally suited to vertical take-off and landing by design. However, they are essentially helicopters having a plurality of rotors in the form of fixed propellers. As such, their reliance on rotary wings results in significant power consumption during both vertical and level flight compared with aircraft equipped with fixed wings. In the context of electrically powered multi-copter UMV (and due to the limitations of existing battery technology), this dramatically reduces the range and endurance of Multi-copters and limits them to relatively short journeys between battery re-charge or replacement.

VTOL (Vertical Take-Off and Landing) Aircraft/Aeroplane

In contrast to the helicopter concept the aeroplane approach is to a powered heavier-than-air aircraft with fixed wings from which it derives most of its lift, at least during the main horizontal transport phase of flight. This allows for rapid forward motion and relatively higher energy efficiency, for example as evidenced by a higher lift to drag ratio. The lift to drag ratio for a helicopter as such is about 1. The VTOL concept providing an initial vertical transport phase in flight to avoid the use of runways and therefore conform to the requirements of urban transport. A number of VTOL aircraft have been developed over the years, mainly for military applications, such as the Boeing V-22 Osprey, the Ling-Temco-Vought XC-142 or the Harrier Jump Jet.

In terms of A-UMV VTOL exemplary current concepts are the Lillium, AeorospaceX Mobi and Airbus Vahana, which like multi-copters have the ability to take-off and land vertically from small foot-print urban landing pads but additionally comprise wings to give lift during horizontal movement. The Lillium concept provides a conventional fuselage with fixed wings but upon the fixed wings are a plurality of rotatable ducted fans which are rotatable between horizontal and vertical orientation, this is supplemented by further ducted fans protruding from the nose region of the aeroplane for giving additional, balancing, vertical lift during take-off. The Vahana concept provides a fuselage with fore and aft aerofoils on which are mounted a plurality of propellers, each aerofoil being rotatable relative to the fuselage to orientate the propellers vertically or horizontally depending on the flight mode.

However, VTOL aircraft also have the ability to transition from vertical to level flight and rely on fixed wings to dramatically improve level flight efficiency reducing power consumption and increasing range and endurance. As a result, particularly in the context of battery powered aircraft, VTOL electric UMV are anticipated to offer a more viable alternative to electric Multicopters for urban transportation over larger distances and/or operate for longer between batteries re-charge or replacement;

The main drawback of VTOL electric A-UMV over their Multi-copter counterparts is the potential complexity of the mechanism(s) required to transition from vertical to level flight and the safety implications associated with a possible failure of such mechanism during a transition.

In summary, the helicopter type aircraft whilst highly manoeuvrable are energy intensive and have limitations in terms of forward flight velocity. The VTOL aircraft overcome this limitation by transitioning to a different geometry for horizontal flight but at the cost of considerable additional complexity. This reduces the potential reliability of the aeroplanes, at the least increases the cost and complexity of maintenance and produces increased regulatory hurdles to get the designs approved.

There is therefore a need to provide an A-UMV which develops the concept of the micro/drone aircraft not only in the use of predominantly electric propulsion was also in the simplicity and potential for mass production, this as opposed to the traditional aircraft development route in a scaled-down format with intrinsic complexity remains, which gives rise to the costs of developing miniaturisation as well as the requirements for sophisticated maintenance.

In addition, all of the above VTOL aircraft whether having reached production or simply concept stage have, many moving parts which are required to move so as to transition from vertical to horizontal flight mode, should any one part failed to transition in synchronisation or simply completely fail to transition then the airworthiness of the aircraft would be severely compromised. There is therefore a need for an A-UMV VTOL capable aircraft of reduced complexity and implied reliability. There is also a need for a A-UMV VTOL capable aircraft capable of failing safe should a transition between horizontal and vertical flight malfunction.

The present invention

The present invention provides an aircraft for use as an airborne, urban mobility vehicle and capable of vertical take-off and landing; the aircraft comprising: a fuselage pivoted between lateral arms of a yoke;

the arms of the yoke extending fore and aft and, at or towards the extremities of the arms: the respective fore portions are linked laterally together by an aerofoil; and the respective aft portions are linked laterally together by an aerofoil;

and at least one of the fore and aft aerofoils having mounted thereon one or more propulsion units.

The present invention in its various aspects is as set out in the appended claims.

Generally, VTOL aircraft rely on mechanisms to either rotate propellers/thrusters with respect to the aircraft wings (as is the case with Lillium or the Harrier and Osprey) or to rotate the entire wings to which propellers/thrusters are attached (as is the case with the XC-142 or the Vahana concept). This calls for relatively complex and potentially highly loaded mechanisms, due to the sheer thrust of the propulsion but also the gyroscopic effect of the rotating propeller or thruster shaft. This can also present a hazard during the transition from vertical to level flight in the event of a mechanical jam but also result in instability as all the forces involved have to carefully be balanced.

The present invention differs from a concept such as the AeroSpaceX VTOL Concept Mobi in that in the AeroSpaceX VTOL Concept a single central wing is pivoted upon a yoke, that wing being rotatable between horizontal and vertical orientations. This means that the wing must also comprise a tail arrangement which adds weight and complexity along with the requirements for elevator controls to change the orientation of the vehicle in horizontal flight and assist with the transition. The present invention provides a simpler arrangement as the fore and aft aerofoils act as both wings and rudder in conjunction with the one or more propulsion units. To this end there are preferably a plurality of propulsion units.

Moreover, the AeroSpaceX VTOL Concept Mobi is expected to require a substantial rotating mechanism as the wing configuration is not believed to offer much moment during transition to assist with the fuselage rotation. For the purposes of the present document a fuselage carries its normal meaning of being the main body of an aircraft and even if not so in terms of size, weight or volume is defined as such by providing the payload carrying function.

For present purposes the yoke comprises all portions of the aircraft upon which the fuselage pivots by means of said pivot.

In the present invention there is preferably provided at least one port and at least one starboard propulsion unit. This enables the aircraft to be steered by altering the degree of proportion exerted by the port and starboard propulsion units, the resulting differential force changing the orientation of the aircraft.

The propulsion units are preferably placed symmetrically upon the aerofoils of the yoke. This reduces the complexity of controlling changes in orientation. The propulsion units are preferably symmetrical fore and aft. This provides a balance due to more even weight distribution between the fore and aft portions of the yoke upon which the fuselage is supported.

More specifically the present invention preferably comprises a distributed electric propulsion system, although in its broadest conception the present invention did not necessarily rely on electric propulsion but may use other conventional propulsion mechanisms, such as a gas turbine. However, electric propulsion is preferred as this provides means for rapid, responsive and directly controllable navigation of the aircraft by means of the propulsion units giving differential levels of proportion.

Distributed electric propulsion architecture may be powered directly by batteries in full electric configuration, by fuel cells or by a hybrid power unit.

An airborne, urban mobility vehicle is an aircraft capable of transporting a human being, or payload of similar weight for a distance and at a height relevant for urban mobility. This AUMV concept of the present invention can easily be scaled from a single seater/rider configuration up to 4-5 seater/rider configuration and beyond. For example, “Present invention-2Rh” refers to a 2 Riders Hybrid Present invention aircraft.

In the present invention the aerofoils are preferably fixed wings in relation to the rest of the yoke and the yoke as a whole only moves with respect to the fuselage at the pivot. This greatly reduces the number of moving parts, providing a simpler and more robust design. This arrangement means that flight control surfaces may not be required and in conjunction with a suitable propulsion unit configuration enables navigation to be undertaken purely by adjusting the output of the propulsion units. For example, the aircraft of the present invention requires no tail section making the design simpler, more cost effective and the mechanical simplicity increases safety as there are fewer parts to potentially malfunction.

In the present invention, the propulsion units are preferably placed fore and aft and further preferably symmetrically and if not literally symmetric then symmetric to the extent of having equal numbers of units, with at least one propulsion units on each aerofoil, this enables manoeuvrability (i.e. navigation) of the aircraft to take place based upon altering the output of the propulsion units.

To this end preferably at least one aerofoil has two propulsion units thereon, the propulsion units being placed respectively port and starboard.

In a preferred embodiment of the present invention four propulsion units on the fore aerofoil and four propulsion units on the aft aerofoil. This provides both the potential for manoeuvrability to be determined entirely by the output of the propulsion units and also provides propulsion unit redundancy so that manoeuvrability may be maintained even if the propulsion unit becomes defective. This greatly increases the safety of the aircraft.

The propulsion units are preferably fixed pitch propeller propulsion units these are simpler and lighter than variable pitch propellers. Variable pitch propellers are not required because in the present invention, particularly with multiple units fore and aft change in unit moment can be significant enough to obviate the need for changing propeller pitch to effect manoeuvrability of the aircraft. The propulsion units are preferably electric propulsion units, this gives a wider range of rotational speed at which both high efficiency and controllability are possible. This is particularly important when manoeuvrability of the aircraft is derived from the propulsion units rather than from control surfaces such as a rudder or ailerons. The use of fixed pitch propeller is made possible by the fact that electric motors operate more efficiently across a wide range of speed and can change speed very quickly, compared with internal combustion engines that are best operated at a constant RPM. With internal combustion engines, the propeller pitch is therefore changed to increase or reduce aircraft speed, whereas with an electric motor the speed of the motor may be changed to also change the speed of the aircraft.

The aircraft of the present invention preferably comprises a flight control unit, the flight control unit controlling power to a distributed electric propulsion system of electric propulsion units driving fixed propellers on all propulsion units. This provides a means to control manoeuvrability in flights of the aircraft based upon differential output from the propulsion units. In some forms of the present invention this provides that:

the flight control unit is configured to manoeuvre the aircraft in one or more of pitch, roll and yaw by means of adjusting the relative propulsive force provided by the propulsion units. Very preferably the flight control unit is configured to manoeuvre the aircraft in one or more of pitch, roll and yaw by means of adjusting the relative propulsive force provided by the propulsion units by means of the relative propulsion moments about the centreline of the aircraft. This is achieved by providing propellers which are paired in CCW rotation and CW rotation. More preferably the flight control unit is configured to manoeuvre the aircraft in one or more of pitch, roll and yaw by means of adjusting the relative propulsion moments (rotational moments, thrust generated moments or a combination of both) about the centreline of the aircraft. Hence, preferably this is why some propulsion units rotate CCW and some CW.

This reduces or preferably obviates the need to auxiliary flight control surfaces, such as rudder, elevators, elevons and ailerons depending upon which selection is made.

For example, Ailerons are normally used to roll the aircraft in level flight (i.e. rotate the aircraft about its centreline, the line defined by the direction of travel). In the illustrated embodiment of present invention (in particular based on Figure 4 propeller configuration) if CW propulsion units 1/2/7/8 are made to rotate faster than CCW propulsion units 3/4/5/6 it creates an imbalance between the moments of the CCW and CW propulsion units and the aircraft rolls to the left (CW moment greater than the CCW moment).

the flight control unit is configured to manoeuvre the aircraft from a vertical take-off to a horizontal flight orientation by means of adjusting the relative propulsive force provided by the fore and aft propulsion units.

the flight control unit is configured to manoeuvre the aircraft in all of pitch, roll and yaw by means of adjusting the relative propulsive force provided by the propulsion units.

In all cases the flight control unit must be fully compliant with regulatory requirements and hence once this had been achieved each additional function serves to improve reliability, reduce complexity and reduce weight as functions normally undertaken by other equipment. For example, a rudder can be omitted as adverse yaw can be accommodated by adjusting the propulsion units (as outlined in principle above for example as described in more detail below). Similarly, ailerons can be omitted as roll (banking) can be accommodated by adjusting the propulsion units. In the same way, elevator and/or elevons can be omitted as pitch can also be accommodated by adjusting the propulsion units.

These features when used all together can mean that, for the purposes of manoeuvring the aircraft in flight, the movable parts of the main body of the aircraft are only the port and starboard pivots of the fuselage and the propulsion units (to the extent that those are in motion to directly produce thrust).

The preferred mode of distributed electric propulsion of the present invention preferably comprises four propeller propulsion units on the fore aerofoil and four propeller propulsion units on the aft aerofoil, the units preferably being placed symmetrically about the fore and aft of the aircraft. DEP (Distributed Electric Propulsion) and specifically DEP in this format enhances lift, reduce drag and hence energy consumption, reduce wing mass/size to offer a reliable, efficient and compact solution to both proportion and navigation/manoeuvrability. Because it allows for smaller wings it reduces drag. The best improvement comes when some (4) of the 8 propellers are also switched off on forward level flight and even greater benefits come from folding the propulsion units that have been turned off. Specifically, the electric propulsion units are required to be individually controllable and this individual control can be naturally extended to control for the purposes of manoeuvring the aircraft. This reduces the number of movable parts. Specifically, the movable parts of the main body of the aircraft are limited to the port and starboard pivots of the fuselage and the propulsion units, the propulsion unit potentially only requiring a rotor and related bearing structures thus potentially giving only n propulsion units plus fuselage as the main moving parts of the aircraft. This greatly simplifies design and production and increases reliability.

Further, the preferred configuration of propulsion units said four propeller propulsion units on the fore aerofoil and four propeller propulsion units on the aft aerofoil gives even greater reliability as up to 50% of the propulsion units may fail while still retaining a reasonable, if emergency, level of manoeuvrability of the aircraft. The invention, such as in the illustrated embodiment, is therefore configured to size each of the 8 propulsion units (such as when driving suitable propellers) such that if a least 2 fail, and indeed if up to 4 fail, the aircraft can still safely land. It may not have the performance to take-off (as to ascend the aircraft needs to accelerate and therefore have a thrust in excess of the mass) but it would be able to land as in this case the aircraft has to be decelerated sufficiently to reach the ground with sufficiently low speed, thus in this case require a thrust level lower than the aircraft mass This is a significant safety advantage of the preferred, illustrated, example of the present invention.

A key option for the present invention is the use of two fixed wings (fore and aft) each equipped with, preferably, fixed propellers/thrusters and instead only the fuselage rotates when transitioning from vertical to level flight as illustrated hereafter:

The invention differentiates over concepts such as MOBI by AerospaceX by the use of 2 wings, fore and aft instead of only one, this has the advantage that the use of 2 wings (forward and aft) allows generating a significant moment to pivot/transition the aircraft from vertical to level flight using differential thrust/lift between the 2 wings/sets of propulsion units without external force. Mobi has only one wing which results in little leverage to pivot the wing. Their outer most propellers are vertically staggered to some extent to offer some moment but they are limited by the fact they only have one wing and so it is not as effective as having 2 horizontally and vertically staggered wings and 2 staggered sets of propulsion units. As such, in order to ease pivoting/transitioning, Mobi likely relies on the mass/inertia of the pod by pulling the wing down via its mechanism and help the wing pivot from vertical to horizontal. This results in significant load the mechanism and render the mechanism critical.

The main benefit of the Present invention UMV is its simplicity over its competitors in this rapidly developing market.

Detailed Description

The present invention will now be illustrated by means of the following figures, in which: Figure 1 - Present invention during Level Flight;

Figure 2 - Present invention in a vertical configuration whilst on the ground;

Figure 3 - Present invention showing rear view with parachute and power module location;

Figure 4 - Propulsion unit (motor/propeller) configuration and labelling;

Figure 5 - Present invention transitioning from Vertical to Level Flight;

Figure 6 - Pitch control of the aircraft in vertical flight;

Figure 7 - Roll control of the aircraft in vertical flight;

Figure 8 - Yaw control of the aircraft in vertical flight;

Figure 9 - Yaw control of the aircraft in level flight;

Figure 10 - Pitch control of the aircraft in level flight;

Figure 11 - Roll control of the aircraft in level flight;

Figure 12 - Redundant Power Distribution and Propulsion System Architecture;

Figure 13 - Alternative wing configuration showing swept and tapered wing;

Figure 14 - Example of control surfaces, “canard” and ailerons;

Figure 15 - Alternative propulsion unit configuration showing a combination of 3-bladed and 2-bladed propellers

Figure 16 - Three tractor motor/propeller version of the Present invention;

Figure 17- Four tractor ducted motor/propeller version of the Present invention;

Figure 18 - Four tractor + four pusher (eight in total) ducted motor/propeller version;

Figure 19 - Present invention fitted with skids for ground support;

Figure 20 - Example of 2-seater version with fuselage acting as tripod and fitted with a pair of nose wheels;

Figure 21 - Example of dissimilar motor technology and configuration;

Figure 22 - Alternative propulsion unit (motor/propeller) configuration and labelling;

Whilst the above figures and the description below describes combinations of features those features may be present separately as defined in the description or in the claims.

The above drawings provide isometric views of the present invention. These drawings illustrate the forward and aft staggered wings, an example of eight distributed electric motor propellers, fuselage capable of housing passenger(s), the yoke with its structure of beams that link the forward and aft wings together as well as the pivot that allows the fuselage to rotate about the yoke assembly.

The Present invention is also depicted in level flight configuration (horizontal or quasihorizontal flight phase during cruise) as well as in vertical flight configuration (vertical or quasivertical flight phase during take-off and landing).

The drawings also provide isometric views of an example configuration of the present invention showing a single passenger aircraft with opened canopy, when the aircraft is on the ground before take-off or after landing.

In flight the canopy is a preferred option for passenger safety and comfort but for clarity the canopy may not always be displayed in some of the illustrations provided.

In the following figures like numerals represent like features. The aircraft 100 of the present invention has the following features:

100, A-UMV

101 to 107 A-UMV variants;

200, passenger fuselage;

202, alternative ‘payload’ fuselage

220, a yoke;

230, 230’, port and starboard arms of the yoke;

240, fore aerofoil;

240’ swept aerofoil example;

242 fore extremity of yoke arm 230 joins to fore aerofoil 240;

250, aft aerofoil;

250’ aerofoil example;

252 aft extremity of yoke arm 230 joins to aft aerofoil;

260, 260’ etc., propulsion unit;

270 Pivot;

280 Canopy;

290 Pilot;

300 Bays for batteries/power-packs

310 parachute bay;

320, 320’ - canards;

330 canard pivot;

340 wing end plates - inward;

340’ wing end plates outward;

350 switched reluctance (SR) motor; and

352 permanent magnet (PM) motor.

Figure 1 shows the present invention during Level Flight; and

Figure 2 shows the present invention in a vertical configuration whilst on the ground.

The ground configuration is essentially the same as the configuration for vertical take-off, it merely being that the optional canopy 270 would be closed on take-off. Similarly, figure 2 shows a pilot/occupant, the present invention is not limited to a passenger carrying aircraft although a preferred embodiment is for passenger carrying. In any case, fuselage 200 comprises a payload carrying space, such as for occupancy by a pilot/passenger(s).

The present invention provides an aircraft for use as an airborne, urban mobility vehicle and capable of vertical take-off and landing; the aircraft comprising:

a fuselage freely pivoted between lateral arms of a yoke;

the yoke extending fore and aft and, at or towards the extremities of the arms:

the respective fore portions are linked laterally together by an aerofoil; and the respective aft portions are linked laterally together by an aerofoil;

Figure 3 shows a rear view (i.e. aft of the aircraft), this illustrates a preferred biplane mode comprising eight propulsion units set upon parallel aerofoils the aerofoils being both horizontally and vertically offset from one another, preferably the aft aerofoil is configured in normal flight to be above the fore aerofoil. This makes the pilot view in line with convention aircraft and enables simpler embarkation and disembarkation. This figure also shows the separate feature of a preferred parachute and/or power module location. This location is preferred as it is readily accessible and is relatively clear of the propulsion units, particularly the propeller blades of propulsion units for the purposes of deploying a parachute and/or replacing a power module.

The manner in which the present invention, as exemplified by this preferred embodiment as shown in figures 1 to 3 operates will now be considered.

As a reference Figure 4 provides a propulsion unit (motor/propeller) configuration and labelling.

Figure 5 - illustrates a key feature of the present invention, specifically the mechanism for transitioning from Vertical to Level Flight. The transition from vertical to level flight is achieved by differential thrust between both fore and aft wings allowing the aircraft to pivot (pitch forward on take-off or backward on landing) and seamlessly transition from vertical to level flight following take-off and with the reverse transition, back to vertical flight prior to landing, as illustrated hereafter figure 5. As can be seen from that Figure the present invention starts out in the configuration shown in figure 2, exerts vertical thrust for vertical take-off and then transitions to the configuration shown in figure 1 configured for horizontal flight.

Rotation of the fuselage relative to the yoke

To accommodate the change in aircraft attitude from vertical to horizontal the fuselage rotates relative to the yoke. As the fuselage mass distribution can be balanced by design, the effort required to level the fuselage is minimal Moreover, as the fuselage does not incorporate any spinning shaft, there is no gyroscopic effect to accommodate, unlike during the rotation of spinning motors/propellers/thrusters. Furthermore, this mechanical arrangement of a pivot is extremely simple and for practical purposes it would be very difficult for it to malfunction in any meaningful way. Even if it did malfunction, there was the aerodynamics of the invention would be non-optimal it would not suggest an immediate disaster situation such as would occur in other designs were multiple components need to simultaneously rotate. Even incomplete rotation of the pivot of the present invention maintains symmetry and hence a higher likelihood of maintaining control.

A set of mechanical stops may be employed to ensure that the fuselage cannot rotate freely about the pivot and is constraint between its level flight position and its vertical flight position. This protects against a failure of the mechanism that would result in a mechanical disconnect (for example a severing of the output shaft of the mechanism).

Hence, a failure of the fuselage during transition has little consequence to the safety of its occupant, Preferably the nose of the fuselage is heavier than the tail as this even avoids the discomfort of possibly flying upside down as the fuselage will always be self-righting.

The rotation of the fuselage pivoted between lateral arms of a yoke is preferably mediated so as to limited or enhance movement that would otherwise occur if the fuselage where freely rotatable with respect the yoke.

The mediation may be by a combination of mechanical stops and a mass and aerodynamic bias. For example, the shape of the fuselage may be aerodynamically designed to result in a moment that would bias the fuselage, under aerodynamic loads, against a first stop designed to prevent over-rotation of the fuselage and keep the fuselage level during cruise (i.e. level flight). Similarly, the weight distribution of the fuselage may be carefully designed to result in a moment that would bias the fuselage, under the effect of gravity, against a second stop designed to keep the fuselage level during vertical flight (i.e. take-off and landing).

The mediation may be by means of a resistive torque such as that provided by a braking arrangement or a clutch.

The mediation may be by means of an actuator to drive rotation about the pivot.

The mediation may be by means of an active control of the fuselage position during level flight (i.e. cruise) to position the fuselage in an optimum position within the air flow so as to minimise the aerodynamic drag of the fuselage and constantly optimise energy efficiency.

Rotation may be limited by mechanical stops, such stops may be repositionable, such as to accommodate different ranges of movement in different flight stages. This protects against dramatic movements, such as flipping of the fuselage due to freak environmental conditions.

Controlling transition of Present invention

As mentioned, the transition from vertical to level flight is achieved by differential thrust between both fore and aft wings allowing the aircraft to pivot (pitch forward on take-off or backward on landing) and seamlessly transition from vertical to level flight following take-off.

Preferred options to effect this transition from vertical to horizontal flight (and by inference in the reverse direction also) are as follows:

Sensors fixed to the yoke, such as on the aerofoils (aka wings) This first option consists in referencing (“fixing”) the flight computer sensors (e.g. compass, gyroscope, accelerometers etc.) relative to the wings of the aircraft. In this case, the flight controller “knows” that it is going to be rotated with respect to the earth referential during the transition, and it is programmed to commands/controls the thrusters to pitch the wings from vertical (e.g. 90deg pitch) to horizontal (e.g. Odeg pitch) whilst commanding the fuselage to remain quasi-level at all time (using any suitable angular/position/attitude sensor). In doing so the wings “lead” by rotating ahead of the fuselage and the fuselage rotating mechanism “follows” the wings and rotate relative to the wings accordingly to keep the passengers in a comfortable level or quasi-level position.

Sensors fixed to the fuselage: A second option consists in referencing (“fixing”) the flight computer sensors (e.g. compass, gyroscope, accelerometers etc.) to the fuselage.

In this configuration the flight computer “does not know” that it is going to be rotated with respect to the earth referential during transition. The fuselage is commanded to rotate which transiently causes it to be slightly out of alignment with the horizontal direction, forcing the flight controller to adjust the thrusters to cause the wings to rotate with respect to the earth referential and level the orientation of the rotating fuselage. In doing so, the fuselage “leads” the wings which are forced to “follow” the fuselage rotation and rotate with respect to the earth differential from vertical to horizontal.

Ideally, in a redundant flight control architecture, both strategies would be implemented, with a redundant set of sensors and computers (fixed to the wings) and a redundant set of sensors and computers (fixed to the fuselage) both in parallel controlling the aircraft attitude.

One flight controller (for example the system fixed to the fuselage, the main system) would be given authority over the other flight controller (for example the system fixed to the wing, the back-up system) and in the event of the main flight controller failing for malfunctioning, the back-up system would safely take over.

By implementing dissimilarity in the software, sensors and computers, this would allow meeting stringent safety requirements, together with the redundant and segregated power distribution architecture and the multitude of redundant thrusters.

Fuselage mechanism

Unlike conventional VTOL aircraft (past, present or in development) that rely on rotating wings and/or thrusters by mechanical means, the proposed concept in fact effect the wings rotation purely via means of differential thrust I differential lift between the forward and aft wings. For passenger comfort, but not required for safe flight, the fuselage is actuated to remain level (e.g. horizontal) or quasi-level. The actuation mechanism of the fuselage is noncritical and is potentially only lightly loaded as both its mass and aero moments can be balanced about its centre of rotation by design. Moreover, in the present invention there are preferably no rotating masses inside the fuselage for the purposes of adjusting flight of the overall aircraft. A mass in this sense being an object intended to navigate the aircraft by altering its aerodynamics (and excluding incidental rotating objects such as gyroscopes, wheels, knobs). The fuselage rotation itself does not generate any (significant) gyroscopic effect, which is often source of instability during the transition of VTOL aircrafts. Unlike known designs the present invention avoids configurations where rotating motors and propellers on wings are moved by mechanisms during transition (e.g. MOBI, Lilium, Vahana, etc.) and as such provides safety and simplicity.

Actuation of the fuselage can be achieved by any suitable mean but may typically be implemented by using:

a) a direct drive rotary actuator, where the output of the rotary actuator is aligned with the axis of rotation of the fuselage; or

b) an indirect rotary actuator, where the output of the rotary actuator is offset from the axis of the axis of rotation of the fuselage and a link and bell-crank connect the rotary actuator to the fuselage axis of rotation; or

c) a linear actuator with its output connected to the fuselage axis of rotation via a bellcrank;

Motor Propeller Sizing Criteria

The motors are sized to ensure that in the event of a failure of either S1 or S2 systems the aircraft may continue to operate albeit under degraded performances particularly during the vertical flight phase during which the aircraft may only be able to land (i.e. control the rate of decent by providing vertical negative acceleration) but not take-off (i.e. provide positive vertical acceleration). Under failure conditions, depending on motor sizing, the motors may have to be over-driven to provide sufficient thrust and may require inspection following an emergency landing.

Ultimately, the aircraft will be equipped with a parachute otherwise referred to a BRS (Ballistic Recovery System) independent from both Normal and Emergency systems as commonly and successfully implemented on light aircraft. This does not preclude to the implementation of a redundant system architecture as parachutes tend to be ineffective at lower altitudes and generally cannot be taken credit from for the purpose of the certification of the aircraft.

AIRCRAFT CONTROL

In both Vertical and Level Flight phases, the Present invention attitude is controlled by combinations of the differential thrust/lift of fore and aft wings propellers and/or the differential thrust/moment of counter-rotating propellers. In both Vertical and Level Flight phases, the Present invention attitude is controlled by combinations of differential moment, from differential rotating speed for CCW and CW propulsion units that allow for the control of yaw and roll (depending on whether the aircraft is flying vertically or level), this is a significant factor in quantitative terms in the present invention and allows for the use of a fixed propeller without disadvantage over a variable pitch propeller.

The information provided for the illustrated, described and preferred aircraft as described herein has been validated by flight in a large indoor enclosed space of scale models (of over 62cm wingspan) of this aircraft and the statements made herein have been validated by flight testing of those models.

Vertical Flight Control

During Vertical flight (e.g. take-off and landing), roll, pitch and yaw are controlled as Shown in Figure 4 which provides an example of motor/propeller configuration and labelling for the present invention and as used in the drawings description.

Figure 6 shows Pitch control of the aircraft in vertical flight. Pitch is controlled by varying the rpm of either front or rear wing propellers, e.g.: if propellers 5, 6, 7, 8 rotate faster than propellers 1,2,3,4 the craft will pitch forward:

Figure7 shows that Roll may be controlled by varying the rpm of either left or starboard wing props, e.g.: if propellers 3,4,7,8 rotate faster than propellers 1,2, 5, 6 the craft will roll to the left:

Figure 8 shows Yaw control of the aircraft in vertical flight. Yaw is controlled by varying the rpm of either CW rotating or CCW rotating propellers, e.g.: if CCW propellers 3, 4, 5, 6 rotate slower than CW propellers 1, 2, 7, 8 the craft will yaw CW:

Level Flight Control

Figure 4 shows an example of motor/propeller configuration and labelling as a reference in the further description.

During Level flight (e.g. cruise), roll, pitch and yaw are controlled as follows:

Figure 11 shows Roll control of the aircraft in level flight. Roll is controlled by varying the rpm of either CW rotating or CCW rotating propellers, e.g.: if CCW propellers 3,4,5,6 rotate slower than CW propellers 1,2,7,8 the craft will yaw CW.

Figure 9 shows Yaw control of the aircraft in level flight. Yaw is controlled by varying the rpm of either left or starboard wing propellers, e.g.: if propellers 3,4,7,8 rotate slower than propellers 1,2,5,6 the craft will roll to the right.

Figure 10 shows Pitch control of the aircraft in level flight. Pitch is controlled by varying the rpm of either front or rear wing propellers, e.g.: if propellers 5, 6,7, 8 rotate faster than propellers 1,2,3,4 the craft will pitch forward.

It should be noted that the motor/propeller configuration and labelling example provided in Figure 4 does not change whether the aircraft is in vertical or level flight.

CCW and CW propellers have a different profile designed to accommodate the direction of rotation whilst providing thrust in the same direction. As such a motor/propeller can only be configured from CCW to CW (and conversely) by physically replacing the CCW propeller for CW propeller (and conversely). It is not simply a case of reversing the motor direction of rotation.

There are however different ways of configuring motor/propellers direction of rotation as illustrated in Figure 22. Figure 22 shows an alternative example of motor/propeller configuration and labelling applicable to the above mechanisms. This alternative configuration is similar to the octocopter motor/propeller configuration commonly implemented on some multi-copters. In this configuration, the direction of motor/propeller 2 and 3 as well as 6 and 7 are inverted compared with the configuration proposed in Figure 4.

System Architecture

The proposed Present invention concept relies on Distributed Electric Propulsion or DEP (in this example 8 electric motors and propellers) to reduce wing surface and drag. This provides additional freedom to implement a redundant system architecture in order to improve reliability and ultimately safety.

The diagram in Figure 12 details an example of redundant architecture comprising of 2 normal systems (S1 and S2) and an emergency system (E):

Figure 12 shows a preferred redundant Power Distribution and Propulsion System Architecture

The layout of each normal systems S1 or S2 is such that they each allow full control of the aircraft. In particular, each S1 and S2 systems is connected to the necessary combination of CCW and CW propellers on each fore and aft wing to allow full pitch, roll and yaw control in both vertical and level flight with either S1 or S2 set of propellers/motors/controllers.

The motors M1 to M8 are in this embodiment distributed evenly between the forward and aft wings and the ESC (electronic speed controllers) required to control each motor may be located inside the wings if space permits, in order to reduce wire count and wire length, or within the fuselage if the wings are too small.

A pair of redundant AP (autopilots) is used to assist in the control of the aircraft in flight and may be located either within the fuselage or wings.

The power source for systems S1 and S2 (e.g. batteries, fuel cells or hybrid units) are located within the fuselage for access but also to better weight distribution of the fuselage in an effort to balance the fuselage mass with its passengers and reduce the loading of the actuation system/mechanism.

An additional Emergency power source (typically an Emergency battery) can be used as a last resort to power either or both Normal systems in the event to a battery failure for example. To this effect, the emergency system is powered by Emergency batteries E1 and E2 (regardless of the normal system power source). To segregate the emergency from the normal system as much as practically possible (e.g. in the event of a fire), the emergency batteries are located in the fore (E1) and aft (E2) wings of the aircraft. In addition, dissimilar battery technology may be implemented for the emergency system, in particular if the normal system is also battery powered. The emergency batteries are size to meet the regulatory requirements for reserve fuel (typically 20 minutes).

Wing design

The wing design of the proposed concept is compatible with any of the modern wing configurations designed to enhance performance.

The wing design of the proposed embodiment may be based on any wing profile, it may consist of a simple straight constant chord wing profile however any other wing configuration may be implemented, for example the wings may be tapered, swept, delta shaped, etc.

Figure 13 shows an alternative example of a swept and tapered wing. The forward wing of the aircraft is known as a swept wing whereas the aft wing is known as a straight tapered (trapezoidal) wing:

The preferred configuration (e.g. figure 1) features 2 backward staggered wings, however a forward staggered arrangement may be implemented and/or the number of wings may be increased to provide additional lift and/or additional control surfaces. For example, additional small wings (sometimes called canard) may be fitted to the hinge point of the fuselage to improve stability and/or act as an elevator.

The current embodiment does not feature any conventional control surfaces, instead relying on differential thrust and/or differential lift to manoeuvre the aircraft about all axis (e.g. yaw, roll, pitch). However, the proposed concept is not limited to this embodiment and conventional control surfaces (e.g. ailerons, elevators, slats, flaps, rudder) may be used to provide additional controllability and/or allow the pilot to retain sufficient control over the aircraft in the event of a loss of power/thrust for example or to reduce the stall speed of the wings.

Figure 14 shows an example of control surfaces, “canard” and ailerons. Additional moveable control surfaces are depicted in the form of “canard” in this example fitted to the hinge of the fuselage as well as an example of more conventional ailerons (322) depicted on the forward wing of this illustration.

Figure 15 illustrates an alternative embodiment of the invention, featuring a combination of 3 bladed propellers (1/4/5/8) and 2 bladed propellers (2/3/6/7). Three bladed propellers are typically less efficient than 2 bladed propellers, however they allow producing very similar thrust with shorter blades than a 2 bladed propeller would produce with longer blades. Due to the reduced blade length of the 3 bladed propellers, the propeller inertia is also reduced, allowing very similar thrust to be produced typically at a higher motor rpm,. For a given propeller pitch, this can allow a 3 bladed propeller to propel the aircraft at a higher forward speed (during level flight) than a larger 2 bladed propellers with the same pitch (forward speed is the product of the propeller pitch and propeller speed). During the slow speed vertical flight phase all propellers have the same performance (i.e. produce the same thrust) but during the high speed level flight phase, where speed matters more than thrust, the 3 bladed propellers may be preferable to the 2 bladed propellers. In this particular example, all four 2 bladed propellers may be stopped and folded during level flight to reduce drag, leaving the faster spinning 3 bladed propellers to propel the aircraft forward at a higher cruising speed, with little change to energy consumption. As an alternative to folding the 2 bladed propellers 2/3/6/7 to reduce drag, these propellers may be stopped in a position parallel to the wing angle of attack (as depicted in Figure 15) to reduce drag, albeit less than if the propellers were folded.

Alternative thruster configurations

The present embodiment features conventional fixed pitch propellers for simplicity and to reduce the weight of the aircraft. In its embodiment the proposed concept features 8 tractor propellers (4 CW and 4 CCW) to provide redundancy and improve safety. Ideally the propellers are distributed along the wing to provide the benefits of what is known as distributed propulsion (whereby blowing on the wing increases lift, allowing for a reduced wing surface and consequently a reduced drag and structure mass).

It is possible, with the advent of modern flight controllers and auto-pilot electronics and software, to configure the aircraft with as few as three propellers (for example two on the aft wing and one of the forward wing). However, the concept should preferably have a minimum of 4 propellers to ensure optimum controllability.

Figure 16 shows a three-tractor motor/propeller version of the present invention. Similarly, the concept may have any number of smaller propellers (for example 6 on each wing for a total of 12 per aircraft), distributed along its wings to provide additional redundancy and further improve the performance of the distributed propulsion. There is also no requirement for an equal number of propellers per wings, and depending on the aerodynamic performance and configuration of each wings, the front wing may have fewer propellers than the aft wing and conversely.

Propellers may be ducted to improve thrust and efficiency (typically by reducing propeller tip losses and via the extra thrust/lift typically generated by the duct itself). The drawback of ducts however is that it may add more mass and complexity to the aircraft. Ducting some or all of the propellers may also protect the fuselage/passengers from the particular risk of propeller blade separation resulting from a failure of the propeller, which is significant because of their rotational speed. This may also be useful in protecting the remaining propellers from being damaged by the failed propeller. Indeed, without the shielding afforded by the propeller duct, the failure of one propeller may result in the failure of all propellers which could be catastrophic

Figure 17 shows a four-tractor motor/propeller version of the Present invention fitted with ducts. The following drawing depicts a four motor/propeller of the Present invention fitted with ducts. An example of structure (in this case 3 struts) required to support is also depicted, illustrating the added complexity that comes with fitting ducts to propellers:

Tractor (i.e. in front of the wings, pulling on the wing) propellers/fans/thrusters may be replaced with pusher (i.e. behind the wing, pushing on the wing) propellers. Alternatively, a combination of tractor and pusher propellers may also be implemented.

Figure 18 shows a four tractor + four pusher (eight in total) ducted motor/propeller version. The following drawings provide a further illustration of the use of ducted propellers, as well as another embodiment of an eight motor/propeller configuration (for redundancy purposes) fitted with four tractor motor/propellers as well as four pusher motor/propellers. In this case, each pair of motor/propeller will typically be fitted with counter-rotating motor-propellers to ensure that each propeller generate thrust in the correct direction.

Generally, more thrust is required in vertical flight phases than in level flight phases. As such, and in the interest of efficiency, some of the motor/propellers may be switched off during level flight phases and possibly fitted with folding propeller arrangements to further reduce drag in level flight.

Similarly, a combination of different types of propellers/thrusters may be implemented to offer a compromise between vertical thrust and level flight speed. For example, a set of large pitch I small diameter propellers may be implemented to allow for fast level flight speed with reduced motor torque, and a set of larger diameter propellers may be implemented to provide high thrust during vertical flight. Both sets of propellers may work together to provide the maximum possible thrust during vertical take-off and landing but the larger diameter propellers may be switched off during level flight to allow the aircraft to travel as fast as possible using as little energy as possible.

In the present invention propellers may be distributed against the wing within the same horizontal plane. It may however be possible to stagger the thrusters of each wing so that some thrusters may be fitted and blow above the wing and other may be fitted and blow under the wing.

Ground stability

On the ground, the aircraft may rest on a conventional landing gear comprising of shock absorbing struts and wheels, however this seems unnecessary for such a VTOL aircraft as it would add weight, complexity and cost to an aircraft that by design should be as light and simple as possible. Instead skids would be preferable, similar to the skids of a helicopter, and could be designed with an element of flexibility in order to dampen the vertical velocity/loads of the aircraft during landing.

Figure 19 shows a further configuration of the present invention fitted with skids for ground support. This version of the present invention features an example of skids to support the aircraft on the ground. This particular version features eight motor/propellers (not depicted) and another smaller aft skid (not depicted) to avoid resting on the aft wing when the aircraft is on the ground:

Figure 20 shows an example of 2-seater aircraft with the fuselage acting as tripod and fitted with a pair of nose wheel. Here, the rotated fuselage forms a tripod with the aft wing and act as a skid I landing gear. A set of wheels or pads may be fitted to the nose of the fuselage (in contact with the ground) and the degree of freedom of the fuselage mechanism may be exploited to provide the compliance and dampening required for a comfortable landing. This may be achieved using a linear spring/damper strut if using an indirect rotary or linear actuation system or a rotary damper with its output directly connected to the fuselage axis of rotation.

Motor configurations

The proposed concept relies on a number of thrusters, at least three but ideally four, to allow controlled and stable flight. Preferably, the proposed embodiment includes eight thrusters arranged in redundant pairs to add an element of safety.

The use of eight thrusters or more, distributed along the wing, enhances lift, allowing for a reduction in wing surface and a consequent reduction in drag and aircraft mass. This is known as distributed propulsion and although conventional engines, as found in existing VTOL aircraft, may be used to power each propeller/fan directly or indirectly (via gearboxes and shafts), distributed propulsion is better suited to electric propulsion, where individual electric motors power, with or without gearboxes, propellers or fans.

The proposed embodiment therefore features eight variable speed electric motors connected to fixed pitch propellers for a simple and practical implementation of distributed electric propulsion.

To further enhance the safety of the aircraft, different motor technologies may be used when redundancy is implemented. For example, in an eight motors/thrusters propulsion unit configuration, four motors may be permanent magnet (PM) rare earth motors (brushed or brushless) whereas the other four motors may be based on a different technology such as switch reluctance (SR) motors (not based on permanent magnets) to avoid common mode failures (for example demagnetisation of the permanent magnets due to excessive temperature) and allow for a safe, dissimilar system. As previously mentioned this illustrated design can fly with only four propulsion units functioning and hence a common failure mode across one type of a set of 4 from the 8 propulsion units leaves the aircraft of the invention in flight. The preferred use of different propulsion unit technologies is in two sets of 4 units per technology.

Figure 21 shows Example of Present invention with dissimilar motor technology. The example depicted in figure 21 feature a combination of permanent magnet (PM) and switched reluctance (SR) motors, with in this particular case PM motors connected to system S2 and SR motors connected to system S1:

Power sources

As discussed previously, although the proposed concept is compatible with traditional fossil fuels (e.g. kerosene, petrol, diesel, gas) engines (e.g. piston engines, turbines), the proposed aircraft is better suited to electric propulsion and therefore require a source of electricity (ideally, but not limited to, high voltage DC if using permanent magnet motors to reduce current and wire gauges).

The source of electricity may be either batteries, fuel cells (for example hydrogen fuel cells) or a hybrid-power unit (for example an internal combustion engine coupled with an electric generator). Or a combination of more than one power source. For example, a hybrid-power unit may provide nominal power and emergency batteries may provide reserve power for nonnominal conditions.

The aircraft would ideally be modular and designed/certified to be compatible with various, interchangeable, sources of power. A dedicated space envelope could be allocated on the aircraft to the power source.

This space envelope I compartment could be fitted with a battery module for a fully electric aircraft with reduced range or customers could elect to purchase a hybrid power module that would fit within the dedicated space envelope and provide enhanced range for customers less concerned with emissions. Alternatively, a fuel cell module may be fitted for a cleaner more readily re-chargeable/re-filled alternative to batteries or hybrid-power unit.

Similarly, in the interest of safety, dissimilar battery technologies may be used between normal and emergency power sources when implementing a fully electric architecture. For example, lighter/smaller Lithium Ion batteries may be used for normal power and NiMH batteries (heavier/bigger) may be used for emergency power. The weight and space envelope penalty of NiMH batteries would be offset by the increased dissimilarity that would improve the safety of the aircraft and prevent common mode failures that may lead to both normal and emergency batteries to fail. An example of common mode failure could be in this case extreme temperatures (hot or cold) that are better tolerated by the NiMH batteries.

In the present invention CW means clockwise and CCW counter clockwise.

Claims,

Claims

1. An aircraft for use as an airborne, urban mobility vehicle and capable of vertical take-off and landing; the aircraft comprising:

a fuselage pivoted between lateral arms of a yoke;

the arms of the yoke extending fore and aft and, at or towards the extremities of the arms:

2. The aircraft of claim 1 wherein the aerofoils are fixed wings in relation to the rest of the yoke and the yoke as a whole only moves with respect to the fuselage at the pivot.

3. The aircraft of claim 1 or claim 2 wherein the propulsion units are placed fore and aft, with at least one propulsion unit on each aerofoil.

4. The aircraft of claim 3 wherein at least one aerofoil has two propulsion units thereon, the propulsion units being placed respectively port and starboard.

5. The aircraft of claim 4 wherein four propeller propulsion units on the fore aerofoil and four propeller propulsion units on the aft aerofoil.

6. The aircraft of claim 4 wherein the propulsion units are placed in equal numbers fore and aft of the aircraft.

7. The aircraft of any preceding claim wherein the aircraft comprises a flight control unit, the flight control unit controlling power to a distributed electric propulsion system of electric propulsion units driving fixed propellers on all propulsion units.

8. The aircraft of claim 7 wherein the flight control unit is configured to manoeuvre the aircraft in one or more of pitch, roll and yaw by means of adjusting the relative propulsive force provided by the propulsion units.

9. The aircraft of claim 8 wherein the flight control unit is configured to manoeuvre the aircraft from a vertical take-off to a horizontal flight orientation by means of adjusting the relative propulsive force provided by the fore and aft propulsion units.

10. The aircraft of claim 8 or claim 9 wherein the flight control unit is configured to manoeuvre the aircraft in all of pitch, roll and yaw by means of adjusting the relative propulsive force provided by the propulsion units.

11. The aircraft of any of claims 8 to 10 wherein the flight control unit is configured to manoeuvre the aircraft in pitch, roll and yaw only by means of adjusting the relative propulsive force provided by the propulsion units.

12. The aircraft of claim 11 wherein, for the purposes of manoeuvring the aircraft in flight, the movable parts of the main body of the aircraft are only the port and starboard pivots of the fuselage and the propulsion units.

13. The aircraft of any preceding claim were the rotation of the fuselage pivoted between lateral arms of a yoke is mediated so as to limit or enhance movement that would otherwise occur if the fuselage where freely rotatable with respect the yoke.

14. The aircraft of claim 13 wherein the mediation is by means of a braking arrangement.

15. The aircraft of claim 13 or claim 14 wherein the mediation is by means of an actuator to drive rotation about the pivot.

Intellectual Property Office

Application No: GB1721837.1 Examiner: Mr Gary Clements

Claims searched: 1-15 Date of search: 25 June 2018

Patents Act 1977: Search Report under Section 17

Documents considered to be relevant:

Category Relevant to claims Identity of document and passage or figure of particular relevance X 1-15 US2014/097290 Al (LENG)-See figures 1-3, and paragraphs 4-22. 56-60 and 64 of description. X,E 1-6 EP3263445 Al (BELL HELICOPTER TEXTRON INC)-See figures la, lb and 2-4p, and paragraphs 32, 33 and 77 of description. A US2018/093765 Al (GRAHAM)-See figure 2. A - US2011/042509 Al (BEVIRT, et al)-See figures 1,2, 11, 19 and 20, and paragraphs 6, 2742 and 53-62 of description.

Categories:

X Document indicating lack of novelty or inventive step A Document indicating technological background and/or state of the art. Y Document indicating lack of inventive step if P Document published on or after the declared priority date but combined with one or more other documents of before the filing date of this invention. same category. & Member of the same patent family E Patent document published on or after, but with priority date earlier than, the filing date of this application.

Field of Search:

Search of GB, EP, WO & US patent documents classified in the following areas of the UKCX :

Worldw ide search of patent documents classified in the following areas of the IPC____________

B64C_____________________________________________________

The follow ing online and other databases have been used in the preparation of this search report WPI, EPODOC, INTERNET.