TECHNICAL FIELD

The present invention relates to a ducted-fan Vertical Take- Off and Landing (VTOL) aircraft in the mini and micro UAV (Unmanned Aerial Vehicle) classes.

BACKGROUND ART

The use of ducted- fan VTOL aircraft in the mini and micro UAV classes for observing areas of territory is known. This type of aircraft normally comprises a toroidal body, which defines the fuselage of the aircraft, supports any surveying and control instruments, is provided with legs or arms at one axial end for standing on the ground and houses a motorized airscrew supported by the toroidal body and mounted to rotate about an axis coaxial with the toroidal body.

In general, this type ducted- fan aircraft is normally preferred to corresponding open-propeller VTOL aircraft due to its greater safety in use, its better propulsion performance and lesser sensitivity to gusty wind.

Attitude control of ducted fan aircraft is normally achieved by arranging a certain number of moveable flaps downstream of the toroidal body' s outlet orifice and connected to the toroidal body.

With regard to flight mechanics, attitude control of the toroidal body in aircraft of the above-described type is performed in the following manner:

Assuming that the ducted-fan aircraft moves in forward flight with a speed VI generally perpendicular to the axis of the toroidal body, and that the apparent wind has a speed V2 equal and opposite to speed VI, the airflow that moves at speed V2 in front of the inlet orifice of the toroidal body is sucked inside the toroidal body by the depression generated by the airscrew and deviated downwards at an angle of approximately 90°. This deviation results in a corresponding change in the motion of the sucked-in airflow, with consequent generation of a first resisting force Dl applied at the centre of thrust (c.p.) of the aircraft, and a first resisting moment MDl, which is applied at the centre of mass or barycentre (C.G.) of the aircraft and has a value given by Dl for the distance between c.p. and C.G. Obviously, Dl is drag, normally applied above C.G. and in the same direction as V2 ; in consequence, MDl is normally a nose- up moment, which tends to move a front portion (in the direction of speed VI) of the toroidal body upwards. In the case where it is wished to exercise attitude control of the toroidal body and at least one of the mentioned flaps is operated for this purpose, the airflow leaving the toroidal body' s outlet orifice is further deviated, with the generation of a second resisting force and a second resisting moment MD2 of the same absolute value and opposite sign to MDl, which are added to Dl and MDl respectively to give a resultant resisting force DR and a resultant resisting moment MDR. Therefore, in each of these quantities, it is possible to distinguish a first value (Dl; MDl), a function of the speed of the ducted- fan aircraft and the conditions of the environment surrounding aircraft, and a second value (D2; MD2) of control, generated by the flaps and complementary to the first with respect to a desired control DR and MDR. To achieve a similar purpose in known ducted fan aircraft, there is a tendency to:

- reduce MDl as much as possible by positioning C.G. as high as possible, so as to reduce the distance between C.G. and c.p. and, consequently, the arm of Dl ;

- place the flaps as far downwards and away from the airscrew as possible, so as to increase the arm of the control forces. These architectural stratagems have the drawback of increasing the height of the ducted- fan aircraft, with a very high position for the barycentre and consequent repercussions on the standing stability of the aircraft on the ground when landing.

Furthermore, the flaps:

are exposed, and therefore subjected to disturbance due to gusts of wind, which has repercussions on the quality of attitude control; and

generate, as mentioned, a force D2 in the same direction as Dl , and therefore for stationary flight in steady wind conditions, a greater inclination of the toroidal body is necessary in order to ensure that the horizontal thrust component of the airscrew equals the sum of Dl and D2.

DISCLOSURE OF INVENTION

The object of the present invention is to provide a ducted- fan VTOL aircraft in the mini and micro UAV classes that enables the above-described drawbacks to be minimized. In particular, the present invention is aimed at improving the operativeness of these aircraft by means of an innovative solution for controlling flight attitude.

According to the .present invention a ducted- fan VTOL aircraft in the mini and micro UAV classes is provided as set forth in claim 1 and, preferably, in any of the following claims directly or indirectly dependent on claim 1.

In particular, according to the present invention, an ducted- fan VTOL aircraft in the mini and micro UAV classes is provided, the aircraft having a centre of pressure or barycentre and a centre of thrust located above the barycentre, and comprising a toroidal body, which has a first axis extending longitudinally along the toroidal body, defines a fuselage of the aircraft and is equipped with means for standing on the ground; and motorized airscrew means housed inside the toroidal body, supported by the toroidal body and mounted to rotate about the first axis; the aircraft being characterized in that the airscrew means comprise a rigid rotor comprising a plurality of blades uniformly distributed around the first axis; each blade having a respective second axis transversal to the first axis, a respective pitch about the respective second axis; and, for each blade, respective first and second control means for respectively collectively and cyclically varying the respective pitch.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will now be described with reference to the attached drawings, which show a non- limitative embodiment, where :

- Figure 1 shows a perspective view of a preferred embodiment of the aircraft according to the present invention;

- Figure 2 shows a partial side elevation, with parts in section and parts removed for clarity, of the aircraft in

Figure 1 ;

- Figure 3 shows a side elevation, on an enlarged scale, of a central portion of the aircraft in Figures 1 and 2; and

- Figure 4 shows, on an enlarged scale and partially in section, a detail of Figure 3 in a different operating configuration.

BEST MODE FOR CARRYING OUT THE INVENTION

With reference to Figure 1, reference numeral 1 indicates, as a whole, a ducted- fan VTOL aircraft belonging to the mini or micro UAV class and comprising a toroidal body 2 with a longitudinal axis 3 and provided at one of its axial ends with feet 4, which are uniformly distributed around the longitudinal axis 3 and project downwards from the toroidal body 2 to allow the toroidal body 2 to stand on the ground. The toroidal body 2 is defined by an inner wall 5 and an outer wall 6, coaxial with the longitudinal axis 3 and side by side. The inner 5 and outer 6 walls are curved with facing concavities and are connected to each other so as to define, between them, an annular duct 7 (Figure 2) coaxial with the longitudinal axis 3 and normally able to contain the survey and control instruments (not shown and comprising, for example, the autopilot) mounted to access the outside of the annular duct 7 through closable access windows 8 made in the outer wall 6.

The aircraft 1 comprises an upper motor 9 that is connected to an upper hub 10, which is connected to the toroidal body 2 in a position coaxial with the longitudinal axis 3 by means of a plurality of spokes 11 extending from the upper hub 10 to an upper inlet lip 12 of the toroidal body.

At its bottom end, the aircraft 1 also comprises a stator 13 (Figure 2) comprising a lower hub 14 coaxial with the longitudinal axis 3 and a plurality of spokes 15, each of which extends from the perimeter of the lower hub 14 to a lower outlet lip 16 of the toroidal body 2 to rigidly connect the lower hub 14 to the toroidal body 2 and support moveable control members 17 (of known type) , which define an adjustable device 17a, of known type, having the function, in use, of controlling the rotation of the aircraft about the longitudinal axis 3.

As is better shown in Figure 3, the motor 9 has an output shaft 18 at the bottom that is coaxial with the longitudinal axis 3 and has an upper portion 19 extending downwards from the upper hub 10 and a lower portion 20, which is integral with the upper portion 19 and has a lower end connected in a rotatable manner to the lower hub 14.

The actuators 22 are arranged around the upper portion 19 of the output shaft 18,. three actuators 22 in this case, which are uniformly distributed around the longitudinal axis 3 and are rigidly connected to the upper hub 10. Each of the actuators 22 is able to operate a respective crank mechanism 23 that, together with the other crank mechanisms 23 and under the thrust of the respective actuators 22, is able to operate an oscillating plate 24 slidingly mounted on the lower portion 20 of the output shaft 18. As is better shown in Figure 4, the oscillating plate 24 comprises an inner portion defined by a sleeve 25, which is coupled to the output shaft 18 by a spherical joint 26 slidingly mounted on the output shaft 18, and through which the lower portion 20 of the output shaft 18 passes.

The oscillating plate 24 further comprises an outer portion defined by a spoked element 28 comprising a central hub 29, an inner surface of which is rotatably coupled, by bearings, to a cylindrical outer surface of the sleeve 25, and a plurality of spokes 30, each of which extends radially outwards from the external perimeter of the hub 29 and is connected at its free end to the free end of the connecting rod of the respective crank mechanism 23. From the above-described assembly, it follows that:

the connection of the spoked element 28 to the upper hub 10 made via the crank mechanisms 23, prevents the spoked element 28 from rotating about the longitudinal axis 3 ;

the coupling of the sleeve 25 to the output shaft 18 via the spherical joint 26 enables the sleeve 25 to oscillate about any axis passing through a centre 27 located on the longitudinal axis 3 ;

the coupling, accomplished by bearings, between the sleeve 25 and the spoked element 28 makes these two components integral with each other with regard to oscillation around the centre 27, i.e. it enables the spoked element 28 to be inclined rigidly together with the sleeve 25 under the thrust of an internal moment applied by the actuators 22, and with regard to sliding along the output shaft 18, i.e. it enables the sleeve 25 and the spoked element 28 to move axially, together with the centre 27, along the longitudinal axis 3 under an internal thrust applied by the actuators 22.

A tubular hub 32 of a rotor 33 is fitted beneath the oscillating plate 24, on the lower portion 20 of the output shaft 18, by means of a radial pin 31, said rotor 33 being rotatably housed within a duct 34 (Figs 1 and 2) defined, coaxially with the longitudinal axis 3, by the inner wall 5 of the toroidal body 2. An articulated transmission 35 is interposed between the sleeve 25 and the tubular hub 32, which defines a torsion- resistant compass able to transmit rotary motion between the output shaft 18 and the sleeve 25. In particular, the articulated transmission 35 is defined by a crank mechanism comprising a lever 36, which is mounted on an external radial appendage 37 of the sleeve 25, having an axis 38 arranged radially with respect to the longitudinal axis 3 and passing through the centre 27. The articulated transmission 35 further comprises a connecting rod 39, a first end of which is connected to the free end of the lever 36 to rotate, with respect to the lever 36, about an axis 40 perpendicular to the longitudinal axis 3 and to axis 38, and a second end of which is connected to the tubular hub 32 to rotate, with respect to the tubular hub 32, about an axis 41 parallel to axis 40.

Therefore, in short, the articulated transmission 35 renders the hub 32 and the sleeve 25 angularly integral about the longitudinal axis 3, enabling, in use, the transmission of rotary motion from the output shaft 18 to the sleeve 25, yet, at the same time, leaving the sleeve 25 free to oscillate around the centre 27 and to slide along the longitudinal axis

3. The rotor 33 comprises a plurality of blades 42 having respective fixed axes 43 extending outwardly from the tubular hub 32 and each provided with both a collective movement and a cyclic movement about the respective axis 43. As is shown in Figure 4, for each blade 42, the rotor 33 comprises a blade-carrier body comprising a pin 44, which is rigidly connected to the hub 32 and extends radially outwards from the tubular hub 32, coaxially with the respective axis 43, which intersects the axes 43 of the other blades 42 at a centre of thrust 45 placed along the longitudinal axis 3 at a distance S (Figure 2) below a barycentre 46 located along the longitudinal axis 3, normally in a position immediately above the upper inlet lip 12. The blade-carrier body of each blade 42 also acts as a hinge for varying the pitch angle of the blade 42 about the respective axis 43 and comprises, for this purpose, a tubular mounting root 47 defined by a generally cylindrical sleeve, which is coaxial with the respective axis 43 and is rotatably fitted on the related pin 44 through the interposition of a bearing system 48, fastened axially with respect to the root

47 by means of a nut 49 screwed onto a threaded end of the respective pin 44. Each root 47 is connected to the oscillating plate 24 via a . respective crank mechanism 50 comprising a crank 51 turning about a respective axis 43, integral with the respective root 47 and extending radially outwards from the root 47, and a connecting rod 52 connecting the free end of the crank 51 to a connecting rod pin 53 projecting radially outwards from the sleeve 25.

In short, the above-described assembly of each blade 42 enables :

rotating blades 42 integrally with the output shaft 18; collectively varying the pitch of the blades 42 about the respective axes 43 through a concordant movement of the connecting rods 52 following the sliding of the oscillating plate 24 along the longitudinal axis 3 (collective control) ; cyclically varying the pitch of each blade 42 about the respective axis 43 through the movement of the respective connecting rod 52 following an inclination of the oscillating plate 24 (cyclic control) .

Since the blades 42 do not have joints for flapping and are not flexible, the above-described rotor 33 is a "completely rigid" type of rotor.

With this solution, collective control is used to control thrust while cyclic control is used for attitude control of the aircraft 1. In fact, collective control simultaneously alters the pitch angle of the blades by the same amount, enabling the thrust to be modified, while cyclic control of the pitch of each blade generates most lift from the blade in the direction in which the pitch angle is maximum and least lift in the opposite direction. As the resultant of the aerodynamic forces acting on the blade is positioned towards the periphery of the blade, toque is generated that enables attitude control of the aircraft.

The use of a ducted-fan VTOL aircraft of small size (mini and micro classes) with a rigid rotor associated with cyclic and collective control of the blades, as in the aircraft 1 of the present invention, brings considerable advantages, which emerge clearly if it considered that the majority of known rotors normally utilized in this type of aircraft are typically fixed and do not have collective movements, and that thrust variation is achieved by varying the number of revolutions, while attitude control is performed via flaps positioned downstream of the annular body, with the drawbacks indicated in the introductory section.

With regard to collective and cyclic control, this is normally used in aircraft with open rotors (not ducted) , such as helicopters. However, in this case, cyclic control is coupled to flapping motion of the blades, or rather an upward and downward oscillating movement obtained by means of hinges or the flexibility of the blades. In these craft, cyclic control operation causes a variation in the sinusoidal lift on each revolution of the blade and the variation in lift causes an upward and downward sinusoidal displacement of the blade on the flapping hinge. Therefore, the plane of the blade is no longer orthogonal to the axis of rotation but is inclined in the direction defined by the given command (orthogonal to the direction of maximum deflection of the blade) . The orthogonal (to the axis of rotation) component of rotor thrust makes the helicopter turn and thus controls the attitude. A rigid rotor, i.e. without flapping motion, would make a helicopter unstable in forward motion as the forward moving blade has a higher speed with respect to the air than the rearward moving one and so torque is generated in forward motion that tends to roll the aircraft over sideways.

The semi-rigid rotor has been derived as a replacement for the rotor in which each blade is provided with flapping movement; the semi-rigid rotor turns as a whole about an axis orthogonal to the axis of rotation or a spherical joint or flexible blades. This type of rotor typically used finds application in aircraft in lower classes and in gyroplanes as it is less complex to construct. Instead, the use of rigid rotors is limited to special applications, such as acrobatic helicopters for example. In fact, in these cases, a flapping rotor would limit the manoeuvres of the helicopters and would not allow performing negative-G manoeuvres, as the blades would deflect downwards and might touch the fuselage of the helicopter. Therefore ^' , for acrobatic types of helicopter, rigid rotors have been studied, in which the flapping movement is less pronounced and is obtained just through the flexibility of the blades. Obviously, these craft require more sophisticated attitude control to avoid roll over in forward motion.

In the case of ducted fan aircraft, the forward motion problem does not exist, as the airflow is always directed along the thrust axis and therefore flapping is not only unnecessary, but is also to be avoided as it would:

reduce or cancel the force differential on the blades and therefore the torque that can be generated;

incline rotor inside the toroidal ring causing interference problems with mechanical parts; and

reduce the propulsive efficiency of the rotor due to the variable gap between rotor and toroidal ring.

Therefore, while "blade rigidity" is a problem in aircraft with non-ducted rotors and equipped with cyclic and collective pitch control of the blades, it represents an innovation and advantage in ducted-rotor aircraft.