TECHNOLOGICAL FIELD

[0001] The presently disclosed subject matter relates to vertical take-off and landing vehicles, and in particular to such vehicles which are additionally suited for use as a roadworthy form of transportation. BACKGROUND

[0002] Surface and road vehicle traffic, especially in large urban areas, is increasing. Even in cities with well-developed public transportation systems which do not interfere with road traffic, e.g., subterranean and/or above-ground mass transit, congestion on roads due to surface vehicles typically results in extended transit times, especially during peak periods throughout the day. [0003] Attempts to develop safe, easy to control and silent electric and hybrid car/airborne vehicle (HCAV) have gained limited success so far. The competing design requirements of the two modes of transportation, including, but not limited to, size, net lift weight, operational range and service height, external dimensions (especially when operating on a surface road), requirements/desire to enable electric-only takeoff and landing, etc., have hindered may such attempts.

SUMMARY

[0004] According to an aspect of the presently disclosed subject matter, there is provided a vehicle for transporting at least one passenger, the vehicle being configured to selectively operate in a driving mode in which it is suited to travel on a road, and in a flying mode in which is it configured to fly in the air, the vehicle comprising a body defining a passenger compartment and carrying a driving arrangement to facilitate the driving and a flying arrangement to facilitate the flying; the flying arrangement comprising two wings and a rotor system, each of the wings comprising an inboard wing articulated to an outboard wing; the rotor system comprising four rotor arms articulated to the body, each having a rotor assembly at a distal end thereof, and a rotor assembly mounted to each of the wings; wherein in the flying mode, the wings and rotor arms extend outwardly from the body, and in the driving mode, the rotor arms are pivoted inwardly toward the body, and the wings are pivoted such that they lie over the body, and folded such that each outboard wing lies over its respective inboard wing. [0005] The vehicle may be configured, in the flying mode, to rotate the wings such that axes of the rotor assemblies mounted thereto are substantially parallel to a yaw axis of the vehicle, thereby being configured for vertical take-off and landing.

[0006] The vehicle may be configured, in the flying mode, to rotate the wings such that axes of the rotor assemblies mounted thereto are substantially parallel to a roll axis of the vehicle, thereby being configured for level flying.

[0007] The vehicle may be configured for short take-off and landing.

[0008] At least some of the rotor arms may comprise elevator flaps.

BRIEF DESCRIPTION OF THE DRAWINGS

[0009] In order to better understand the subject matter that is disclosed herein and to exemplify how it may be carried out in practice, embodiments will now be described, by way of non-limiting example only, with reference to the accompanying drawings, in which:

[0010] Figs. 1A and IB are perspective views of a vehicle in a driving mode and a flying mode, respectively;

[0011] Fig. 1C is a perspective view of the vehicle in a hovering mode; [0012] Fig. ID is a perspective view of the vehicle in which wings thereof are shown folded;

[0013] Fig. 2A is a top view of an outboard folding mechanism of the vehicle illustrated in Figs. 1A and IB;

[0014] Fig. 2B is a side view looking outwardly of the outboard folding mechanism;

[0015] Fig. 2C is a front view of the outboard folding mechanism; [0016] Figs. 3A and 3B are top views of a tilt-pivot mechanism of the vehicle illustrated in

Figs. 1A and IB;

[0017] Figs. 4A and 4B illustrate first and second steps of a wing-pivot sequence;

[0018] Fig. 5 illustrates a wing-tilt sequence;

[0019] Figs. 6A and 6B illustrate increased root cross-section of a wing of the vehicle; [0020] Figs. 7A and 7B illustrated square tube reinforcements in the wing spars of the vehicle; [0021] Figs. 8A and 8B illustrate different positions during transition between the flying and driving modes of the vehicle;

[0022] Fig. 9 is a top view of the vehicle;

[0023] Fig. 10 illustrates a bumper of the vehicle; and [0024] Fig. 11 illustrates an algorithm for determining operational parameters of the vehicle during use.

DETAILED DESCRIPTION

[0025] As illustrated in Figs. 1A and IB, there is provided a vehicle 10, which may be selectively transformed between a driving mode and a flying mode. In its driving mode, the vehicle 10 may be configured to operate on a surface while complying with relevant regulations governing roadworthiness, for example in connection with one or more of its dimensions, configuration, passenger capacity, weight, etc. In flying mode, the vehicle may be configured to hover, fly, and take off vertically and/or using a short runway, which are referred to in the art and in the present description as vertical take-off and landing (VTOL) and short take-off and landing (STOL), respectively.

[0026] The vehicle 10 defines a roll axis, a pitch axis, and a yaw axis, corresponding to aircraft principal axes, and which are indicated in Figs. 1A and IB as X, Y, and Z, respectively.

[0027] The vehicle 10 comprises a body 12 defining a passenger cabin therewithin configured to accommodate at least one passenger (for example including a pilot). It further comprises a driving arrangement 14 configured to facilitate operation in the driving mode, and a flying arrangement 16 configured to facilitate operation in the flying mode. It will be appreciated that some components of the driving arrangement 14 may be used during the flying mode, and vice versa.

[0028] The vehicle 10 may further comprise a controller configured to direct operation thereof. It will be appreciated that while herein the term “controller” is used with reference to a single element, it may comprise a combination of elements, which may or may not be in physical proximity to one another, without departing from the scope of the presently disclosed subject matter, mutatis mutandis. In addition, disclosure herein of the controller carrying out, being configured to carry out, or other similar language, implicitly includes other elements of the vehicle carrying out, being configured to carry out, etc., those functions, without departing from the scope of the presently disclosed subject matter, mutatis mutandis.

DRIVING ARRANGEMENT

[0029] The driving arrangement 14 may comprise a plurality of wheels 18 configured to propel the vehicle along a surface road. According to some examples, each of the wheels 18 is mounted to a wheel support 20 extending outwardly from the body of the vehicle. The distances between the wheels, e.g., along the pitch and/or roll axes of the vehicle may be such that they do not exceed dimensions allowed for a roadworthy vehicle.

[0030] The driving arrangement 14 may further comprise other elements (not illustrated) configured to facilitate operation on a road surface, including, but not limited to, brakes, a steering system, etc.

FL YING ARRANGEMENT

[0031] The flying arrangement 16 may comprise a wing system and a rotor system.

Wing System [0032] The wing system comprises two wings 22, each extending outwardly, at least in the flying mode of the vehicle, from the body generally along the pitch axis.

[0033] As best seen in Fig. IB, each wing 22 comprises an inboard wing 22a articulated at a proximal end thereof to the body, and an outboard wing 22b articulated to a distal end of its respective inboard wing. Each inboard wing 22a may be articulated to its respective outboard wing 22b by one or more folding mechanisms, for example as described below.

[0034] In the driving mode of the vehicle, the inboard wing 22a has the ability to pivot backward, for example by about 90° (for example as seen in Fig. 1A). In the flying mode, the inboard wing 22a can be in one of two positions, one vertical wing configuration for the hovering mode (for example as seen in Fig. 1C) and one horizontal wing for level flight (for example as seen in Fig. IB).

Folding Mechanism

[0035] Before the inboard wings 22a are pivoted, the outboard wings 22b may be folded, each to overlie its respective inboard wing, for example as illustrated in Fig. ID. Accordingly, a folding mechanism, generally indicated at 24, is provided to facilitate the folding. According to some examples, e.g., as illustrated in Figs. 2A-2C, each folding mechanism 24 comprises two linear actuators 26, one on each side of the wing spar 38. In the flying mode, a hinge 28 connecting the inboard and outboard wings 22a, 22b is automatically locked, for example with two locking pins 30 driven by a rotary actuator 32. A triangle-shaped bell-crank 34 may be provided to connect the two actuators 26 and the hinge pin 30 to allow the required range of movement. The hinge plates are attached to the last (i.e., distal) rib of the inboard wing 22a and the first (i.e., proximal) rib of the outboard wing 22b. An upper pivot 36a of the bell-crank 34 holds the hinge pin 28a around which the outboard wing 22b turns. Two lower pivots 36b of the bell-crank 34 are connected to the linear actuators 26. During folding the locking pins 30 are automatically disengaged, and the inboard actuator pushes the bell-crank outwardly for example by 90°. At the same time, the outboard actuator pushes the bell-crank to complete the fold, for example 180°.

[0036] The outboard wing folding feature is secured by the disposition of the locking pins 30 close to the lower skin through the spar 38 of the inboard wing 22a and the spar 38 of the outboard wing 22b that overlaps with the spar 38 of the inboard wing with lugs. The two locking pins 30 keep the inboard and outboard wings 22a, 22b locked for flight. The added pin mechanisms restore the wing bending load path.

[0037] The shear load path web of the spars/skin is broken down at the fold mechanisms. The bending moment distribution is decoupled into axial tension and compression and transferred through the pins. Since the spars of the wing are reinforced with composite tubing, these spars are equipped to take more axial loads than the traditional spar of a typical wing. The direction properties of the composites are leveraged at this point to provide specific directional strength to these spars.

Tilt-Pivot Mechanism

[0038] As illustrated in Figs. 3A and 3B, each wing may further comprise a tilt-pivot mechanism 50, configured to facilitate selectively pivoting or tilting the wings. Fig. 3A is the top view of the right-hand-side of the mechanism assembly. The tilt-pivot mechanism 50 allows motion in two directions while transferring loads from the wing rigidly and efficiently, functioning as the critical load path between the wing, wing rotor blades, and the fuselage. The wing must be pivoted to allow the vehicle to be driven on the road. [0039] The tilt-pivot mechanism 50 comprises two forward and aft carry-through members 52a, 52b (collectively indicated by 52), two hinges 54, a wing-tilt linear actuator 56, a wing- pivot actuator 58, and first and second fairing-fold actuators 60a, 60b. The forward carry- through member 52a connects the forward spars 38 of the wings 22a, 22b with a hinge 54 between them, which allows the pivot of the wing. The hinge axis is slightly tilted from the yaw axis. This allows the wings 22a, 22b to overlap with each other when they’re pivoted, without any additional mechanism. Thrust bearings are inserted between lug plates of the hinge 54. [0040] The aft carry -through member 52b is connected to the aft spars 38 of the wings, with a joint that can be disengaged and disconnected when folding the wing. This second carry - through member sits on ball bearings 62, which allows the tilting of the wing.

[0041] The inboard tilting engages the two carry -through members 52 while in the flight and hover positions. The carry-through members 52 may connecting the left and right wings using I-beam cross-sections. The cross-section of these I-Beams is tailored so that it offers the specific strength to support the wing-up bending loads. This tailoring of the I-Beam cross-section is geared toward reducing weight and yet has the strength/stiffness properties of a traditional wing- box structure. While in the vertical-flight position, these members maintain the continuous load path by positioning the forward member about an axis of the tilt bearing. The two thrust bearings are installed in a front-to-front configuration in order to balance the shear loads from the wing and flow it down to the fuselage forward bulkhead. This unique use of the thrust bearing allows for the folding of the wing at the inboard end yet maintains the shear load path.

[0042] The location and separation of the carry-through members 52 may minimize the bending moment at the folding hinge while remaining within the street-legal dimensions, particularly the width constraint, minimizing the wing pitching moment transferred to the carry- through members. The further outboard the hinge moves, the further forward the hinge needs to be, in order to remain within the width constraint.

[0043] Two linear actuators are connected to the forward carry-through member. The extension of these actuators tilts the wing, for example 90°. According to some examples, one or both comprises a ball-screw electric actuator. They are configured to transfer large tensile/compressive loads without having electric power to them, making them part of the load path. [0044] The wing -tilt linear actuator 56 and wing -pivot actuator 58 are installed near the hinge 52 at the forward carry-through member from the wing forward spar. These actuators extend and shrink to pivot the wing on top of the fuselage. When the wings are extended, these actuators brace the moment around the yaw axis, which can be large particularly during the vertical flight. [0045] The tilt-pivot mechanism is configured to transfer the aerodynamic and thrust loads efficiently to the fuselage, and also it must secure the wing securely with no excessive movement. Therefore, it may be configured to provide multiple constraints and load paths efficiently, and to constrain all degrees of freedom from the wing by multiple means. The rigid wing root may mitigate vibration and flutter, to which high aspect-ratio wings may be prone. Providing multiple constraints and load-paths may also facilitate preventing catastrophic failure caused by a single point of failure. Such measures to constraint shear forces and moments in all three axes are shown in Fig. 3B.

[0046] It is equally applicable to maintain the direction of the thrust vector from the wing rotor blades constant as much as possible so that the generated thrust is effectively used to lift the vehicle and the flight control system is not compromised.

[0047] Except for the hinge pin which requires a high-strength durable material such as steel or titanium due to the high shear load, this design allows the use of aluminum for the majority of the section. This may reduce manufacturing costs and/or weight.

[0048] As illustrated in Figs. 4 A and 4B, when pivoting the wing from the flight mode to the stowed drive mode, the fairing moves up from the hinge near the forward carry -through member first, pushed by a micro linear actuator. The aft spar lock pins are then disengaged to allow the movement of the wing. Finally, the wing -pivot linear actuator extends to push the wing backward.

[0049] As illustrated in Fig. 5, prior to take-off and transition, when the wing tilts upward as described above, the aft fairing is pushed up, for example by 90° to clear the fuselage roof. Then the linear actuators on the FWD carry-through member extend to push the carry-through member so that the wing tilts around the bearing on the AFT carry -through member, for example 90° Wing Structural Design

[0050] The wings may comprise a truss-stiffened wing spar 38, thereby providing stiffness, which can override the effect of the mass increase. With this feature, the bending and torsional mode frequencies will go up and flutter will occur at higher speeds. For example, each of the wings may comprise a spar section 38 configured to provide adequate strength and free specific degrees of freedom to allow for the folds described above. These features not only provide the required load path for the lifting but also reduces the overall weight of the aircraft by taking advantage of the sectional properties in that configuration.

[0051] The root cross-section area may be provided such that the second moment of inertia at the root is large enough to reduce the deflection of the wing, for example as illustrated in Figs. 6A and 6B. Another favorable effect of this is the minimization of the interference drag at the wing root. The fuselage shape has been aerodynamically optimized to produce as much lift as possible while minimizing the drag. Throughout the analysis, it was shown that the larger wing cross-section area, i.e., larger fairing at the wing -fuselage joint, had a favorable effect on reducing drag and increasing lift.

[0052] In order to increase the bending rigidity (El) of the wing, the conventional I-section spars may be reinforced with a composite square tube, which runs from the root to the tip, for example as illustrated in Figs. 7A and 7B. One or both of the forward and aft spars may be reinforced with square tubes 80, making them integral parts of the spar web and the upper skin of the wing. Tubings with a smaller cross-section run alongside the aft spar. The forward and aft spars are cross-linked with a square tube, which creates a truss system to transfer the loads between the forward spar and the aft spar at the root. In the root section where the cross-sectional area increases, square composite tube diagonal members enable the load transfer between the lower spar cap to the stiffer upper cap section. They efficiently re-route the loads to the forward spar upper cap composite tube without a significant weight increase.

[0053] Such square composite tubes have suitable mechanical properties, particularly for tension and compression, flexibility of the design, and readily available commercially. These advantages make them ideal for the integral reinforcement of the wings with a unique planform. [0054] The aeroelastic performance of the wing may be optimized by utilizing the directional properties of the composite materials in order to avoid divergence. Finite Element Analysis (FEA) may be used to optimize the structure of the wing. Wing Planform

[0055] According to some examples, the wings may be characterized as follows:

• Wingspan: 15.7 m (619")

• Wing area: 16.4 m ² (176.7 sq ft)

• Mean Aerodynamic Chord: 983 mm (38.7")

• Aspect Ratio: 15.1

[0056] The wing 22 may have a relatively narrow chord and a high aspect ratio. Additionally, for example as seen in Fig. 9, the leading and trailing edges may each be broken into multiple sections, e.g., by providing the wings with a kink 82 therebetween. This so-called “kinked” planform may increase the wing area while foldable for the drive with the rotor mounted on the leading edge and to clear the arm hinges on top of the fuselage.

[0057] Another benefit of the kinked planform is that the aerodynamic center of the wing is shifted aft. The wing root must be closer to the nose than typically desired for fixed-wing aircraft, in order to have a sufficient wingspan. The kink helps the aerodynamic center to stay in the desired location.

[0058] The maximum chord length may be limited as the wings must fold on top of the vehicle, and the area of the wing that can overlap must be kept minimal as it will affect the design of the folding hinges. However, this requires the vehicle to have an aerodynamically efficient wing planform with a high aspect ratio. The amount of induced drag is inversely proportional to the aspect ratio of the wing.

Rotor System

[0059] Reverting to Figs. 1 A through ID, the rotor system comprises two forward rotor arms 90 and two aft rotor arms 92, and a rotor assembly 94 mounted on each. It further comprises a rotor assembly 94 mounted on each of the wings. Each of the rotor assembly comprises a rotor hub and a plurality of rotor blades attached thereto. The rotor blades may be rigidly attached to their respective rotor hubs.

[0060] During vertical take-off and landing, the wings may be tilted about 90° around the pitch axis, for example as described above, such that each of the rotor assemblies mounted thereon is positioned so as to rotate about an axis substantially parallel to the yaw axis. According to some examples, the rotor assemblies may be configured to each rotate in an alternate sense as an adjacent rotor assembly, e.g., the rotor assemblies on the right wing and on the left rotor arms may rotate in a clockwise direction, while the rotor assemblies on the left wing and on the right rotor arms rotate in a counter-clockwise direction.

[0061] During level flight, wings may be tilted such that each of the rotor assemblies mounted thereon is positioned so as to rotate about an axis substantially parallel to the roll axis. Moreover, during level flight the rotor assemblies on the rotor arms may be deactivated.

Wing Rotors

[0062] The tilt-wing configuration has a positive aerodynamic effect. The rotation of the wing-rotors is set to counteract the wingtip vortex and reduces the induced drag. Additionally, the accelerated flow in the disturbed mass of air pushed aft by the propeller (propwash) maintains its direction with respect to the wing constant throughout the flight. Therefore, the section of the wing in the propwash consistently receives the energized flow at the optimal angle of attack regardless of the airspeed and the angle of attack of the vehicle. The wing generates additional lift while the increase in the drag is kept low; the increase in parasitic drag from other parts of the vehicle such as the fuselage is avoided. During the transition when the angle of attack of the wing can be very high, the wing-rotors can help delay the stall as the effective angle of attack can remain low due to the direction of the propwash being constant.

[0063] Another advantage of the tilt-wing concept is that it disposes the wing so as to minimizing the blockage. The smaller distance between the rotor blades and the wing leading edge also has benefits, as the wing benefits from the energized flow in the propwash before it dissipates.

Lift rotor blades on the arms (forward and aft rotor arms):

[0064] Two lift rotor blades are installed on the forward arm and two additional lift rotor blades are installed on the aft arm. [0065] Limiting the forward thrust operation of the rotor blades only to the wing rotor blades may increase the airflow velocity over the wing, helping improve the lift. Another advantage is the motor efficiency, it is more energy-efficient to operate two motors in a medium load operation point instead of operating six motors in a low load operation point. Lastly, the four static lift rotor blades are able to be optimized for lift efficiency, further improving its disk loading and power loading factor. [0066] In order to convert the vehicle from its flying mode to its driving mode, the aft rotor arms 92 are pivoted rearwardly, for example as illustrated in Fig. 8 A. The outboard wings 22b are then folded to overlie the inboard wing, for example as illustrated in Fig. ID. The inboard wings 22a, carrying the outboard wings, 22b, are then pivoted rearwardly to overlie the body, for example as illustrated in Fig. 8B. The forward rotor arms are then pivoted rearwardly to overlie the body, for example as illustrated in Fig. 1 A.

[0067] In order to fully turn and store the wing above the fuselage, the midsection of the vehicle may comprise a mechanism configured to automatically remove a part of the fuselage to make room for the wing (for example as illustrated in Fig 4 A). It also includes a latch to securely close the fuselage during flight and all electrical systems necessary to operate. Additionally, several different sensors will be required for the safe operation of such mechanisms. This includes, but is not limited to proximity sensors, IR sensors, ultrasonic sensors, and limit switches.

Rotor Arms

[0068] The stiffness of the rotor arms may limit the design more than does their strength. Accordingly, a truss-type configuration may be provided. According to some examples, each rotor arm comprises two composite square tubes of dimensions 107x107mm and thickness of 3 mm. The plies in the tubing system are selected to take the majority of the tension-compression load couple resulting from the bending load produced by the tip rotor. The combination has excellent bending performance yet with the minimum weight possible. The arms are rigidly secured at the base to the fuselage to carry the bending moments into the aft bulkhead of the fuselage.

VTOLANDSTOL

[0069] The vehicle is designed as a VTOL drive and fly vehicle, but is also capable of short take-off and landing. In order to shorten the runway length required, the vehicle is configured to utilize the tilting wing and the driving mechanism during different flight phases.

[0070] The tilting of the wing divides the thrust into a vertical force (i.e., parallel to the yaw- axis) and a horizontal force (i.e., parallel to the roll-axis). By this, enough speed can be provided to ensure the lift-off with the wing and at the same time minimize the runway length by enhancing the lift with the vectored thrust of the rotor blade. This allows us to utilize the driving capabilities of the vehicle during STOL, and reduce the duration of axial flight where the high propulsive power demand exists. In this case, the wheel drive system is accelerating the vehicle up to its maximum driving speed and then the rotor blades take over. This includes a system configured to facilitate switching between the wheel drive system and free-spinning wheels during the acceleration with the rotor blades, for example comprising a reverse-centrifugal clutch which disengages from the wheels at a maximum driving speed.

[0071] The vehicle may be configured to utilize the in- wheel motor and the wing to perform a short take-off, thereby saving energy.

[0072] The rotor assemblies on the rotor arms may be in a standby mode during the conventional takeoff, thereby allowing flight control can secure safety by stabilizing vehicle attitude in case of critical failures, for example, a proposition rotor failure or for an in-wheel motor failure during takeoff. Thus energy reserves for failure cases can be further optimized compared to other conventional aircraft.

[0073] The vertical take-off and the runway take-off consume a similar amount of energy in the initial phase. During vertical take-off, the vehicle must be accelerated in the vertical direction and during a runway take-off, the car has to be accelerated in the horizontal direction up to rotation speed. In addition, in VTOL mode, upon reaching 66 ft (20m) altitude, the vehicle has to be accelerated horizontally until it reaches minimum climb speed. This phase is called transition. During the transition, the tiltable rotor blade systems are rotating around their pitch axis. This is a critical phase since the thrust used to accelerate the car in the vertical direction is therefore redirected in the horizontal direction. As a requirement the vehicle is not supposed to dip, meaning losing altitude, during that phase. In conclusion, this phase is a costly endeavor in terms of energy. This phase is only necessary during vertical take-off and landing respectively. [0074] Using this method allows the vehicle to provide enough thrust to vertical take-off and accelerate to and hold the necessary relative velocity to provide sufficient aerodynamic lift and at the same time keep the drag to a minimum during level flight. In an unexpected and rare case of an emergency, the lift rotors can be utilized to lower or stop a rapid descent and are therefore part of the safety measures. STABILITY AND CONTROL

[0075] As illustrated in Fig. 9, the vehicle may further comprise ailerons 96 on the wings, a rudder 98 on the vertical fin, and/or elevator flaps 100 in the arms.

[0076] The forward and aft arms may be profiled and placed to provide additional lift and work alongside the other surfaces to provide the stability and controllability to the vehicle.

[0077] The rotor arms may function as the horizontal stabilizers and canards. The long moment arms from the center of gravity make them very effective in controlling the vehicle. Particularly, the elevator flaps as part of the canard and the horizontal stabilizer help reduce the elevator surface area and the required deflection, which in turn reduces the trim drag. The elevator authority is also increased because of the configuration, widening the allowable CG range for the forward flight. The increased elevator authority is also critical for the short takeoff and landing (STOL) operation and transition when the airspeed is low.

[0078] For lateral and directional control, the wing has conventional ailerons and the empennage includes a vertical fin with a rudder. In order to keep the hinge moment and the actuator weight, additional drag and adverse yaw low, and to reduce the required aileron deflection to minimize the risk of a tip stall, the ailerons benefit from the high aspect ratio of the wing and the long moment arm. The vertical fin can be made into a 2 V-shaped surface configuration as well, as necessary. However, the vertical fin can also be made as a single surface, depending on the [0079] The retractable arms, the kinked wing planform, and the placement of the rotor blades contribute to the reduced control and stability effort, which in turn, reduces the power requirement as the trim drag during the forward flight or trim thrust during vertical flight. [0080] Proper center of gravity management is also critical for a safe and efficient transition, however, if necessary, the allowable CG range for the vertical flight can be shifted by adjusting the locations of the forward and aft rotors which are on the foldable arms. In other words, adjusting the opening of those arms. Although not as fast as the movement of the vertical thrust centroid, this shifts the aerodynamic center in the same direction and can be used to adjust the elevator sensitivity as well, since they are also part of these arms.

[0081] During the vertical flight, the six rotor assemblies are almost equidistant from the CG, which allows for a stable flight and straightforward flight control. The empty-weight CG is near the aft-limit, and loading passengers and payload in the cabin shifts the CG forward. The wing, control surfaces, and rotors were placed so that the vehicle performs optimally in this range. [0082] As illustrated in Fig. 10, one or more bumpers 110 may be provided in order to ensure that mechanical impacts which may occur, for example when the vehicle is operating in its driving mode, do not affect the airworthiness of the vehicle when in its flying mode, portions of the vehicle may be configured to properly absorb such impacts. Accordingly, relevant areas of the vehicle, that are mostly expected to be involved in such an accident, may be protected by a shock absorbing system. The bumpers may be made of absorbing material, e.g., foam panel. Panels may be attached over the outer skin of the vehicle, such as the left and right sides of the rudder. A bumper fin surface may be provided to protect the vehicle from a collision from behind. Bumpers may further be provided on other vehicle components, including, but not limited to, the rotor blades when they are folded. They are intended to hit any object from behind first, without damaging the rest of the vehicle, and be easily replaceable. These surfaces also provide additional longitudinal stability to the vehicle. The incidence of this surface can be adjusted to match the center of gravity, if necessary.

PROPULSION OPTIONS

[0083] According to some examples, the vehicle may be designed to operate using electrical power, for example without a heat engine such as an internal combustion engine. Accordingly, it may comprise one or more batteries, fuel cells, or any other similar components. [0084] According to other examples, the vehicle uses hybrid power, e.g., comprising one or more engines, one or more batteries, an electrical generator, and a power distribution system. Engines may include turboshaft engines, aviation and/or conventional gasoline/diesel piston engines, and/or hydrogen fuel cells.

[0085] The power distribution system is intended to distribute the power to the flight systems, drive systems, rotors, flight control surfaces, lights and any other electrical source. The distribution system includes the capabilities to prioritize and redistribute the power of the battery as well as the engine. It is designed to foremost secure safe flight for the passengers and/or payload. [0086] The system also provides double redundancy by design. In the very unlikely case of a power loss of both propulsion systems the vehicle is still able to glide to the next landing opportunity and perform a conventional landing.

[0087] To reduce the noise in crowded areas the vehicle may be configured to take off using power from batteries only.

CALCULA TING OPERA TIONAL PARAMETERS

[0088] As illustrated in Fig. 11, the controller may be configured to calculate target operational parameters, such as target rotational speeds for the rotors, for example based on physical and operational parameters of the vehicle, in order to achieve a predetermined rate of acceleration. It may be further configured to direct the vehicle to operate per the calculated target operational parameters.

[0089] Input data, including, but not limited to, overall motor parameters (e.g., the energy of the battery, the maximum torque, etc.), rotor assembly parameters (e.g., as given by the inertia of the fan, the maximum RPM, the thrust, etc.), in- wheel motor parameters, design parameters (maximum take-off weight, dimensions, etc.), and aerodynamic parameters (lift coefficients, drag coefficients, etc.).

[0090] The thrust, the aerodynamic forces as well as the friction between the wheels and the ground are taken to calculate a force vector. In this case the mass maybe set to the maximum take-off weight. With this information the controller may calculate an acceleration vector and a velocity vector, as well as the horizontal and vertical distances traveled. If, e.g., the acceleration is higher than the maximum allowed, the RPM may be adjusted until this criteria is satisfied. After that the energy demand is calculated and the distance of this time frame is added to the runway length and the next time step is calculated.