DESCRIPTION The present invention relates to the field of aeronautical engineering and more specifically the field of the remotely piloted aircrafts, with propeller propulsion, able to perform vertical take-off and landing and then which do not need take-off/landing runway, catapult, sling or hand throwing. In particular, the invention relates to the subclass of the aircrafts called "tail-sitter" in technical language, that is the set of aircrafts performing the take-off with wings, fuselage (if existing) and propelling plant in vertical attitude and with thrust directed upwards, they subsequently perform a transition/rotation to pass to the advance flight and then they place in horizontal attitude (as a conventional airplane) to perform cruising.

Currently the aircrafts of "tail-sitter" or "tilt-body' type are practically absent on the market as they are a little efficient, relatively heavy, and more complex from the point of view of the automatic checks they need, along the whole envelope.

The VTOL "Tail-Sitter" aircrafts described in the state of art suffer from low autonomies in hovering, due to the reduced size of the rotors, consequently this problem makes them machines which take-off vertically, but which do not fly economically as helicopters. The size of the rotors cannot be increased arbitrarily as it causes increase in the aircraft weight and a shifting of the centre of gravity towards the wing leading edge, with consequent worsening of stability and longitudinal controllability.

Each inherently stable aircraft must have (necessary condition) positive zero lift moment coefficient and negative pitching stiffness. These two requirements are opposite to the property of an isolated wing developing positive lift coefficient along the whole span under normal cruising conditions, then in order to manage the longitudinal stability of the all-span aircraft able to take-off vertically one has to adopt a wing plan far from the optimum one guaranteeing the elliptical circulation distribution along the span. Furthermore, one has to warp the wing (aerodynamic and/or geometric warping) and to plan part thereof so as to develop negative lift to equilibrate the pitching moments. A little stable, little manoeuvrable, aerodynamically little efficient aircraft results.

The low aerodynamic efficiency which is obtained to guarantee stability and longitudinal controllability could be compensated with the increase in the wingspan (or in the wing lengthening). However, generally the aircraft weight is more or less proportional to the third lengthening power, then a greater aerodynamic efficiency would be obtained to the detriment of the weight.

Furthermore, the requested wing surface being equal, an aircraft with a high lengthening suffers from the effects due to the decrease in the number of Reynolds: the wing profiles do not carry as they should do, they resist more, they have irregular polars, discontinuous at certain incidences. The high wing lengthening, substantially, is useful to reduce the induced resistance, but generally it penalizes the parasitic resistance and the aircraft weight.

If, in order to avoid the decrease in the number of Reynolds, the wing surface is increased, the span being equal, a weight increase is however obtained and consequently the lengthening is reduced, with increase in the induced resistance. A "Tail-Sitter" aircraft has unsurmountable limits linked to the particular all-wing configuration, consequently it is impossible increasing the lightness, stability/control and efficiency thereof at the same time. Practically, upon working on a "Tail-Sitter", it is not possible to bring the performances thereof to the limit, neither in the horizontal flight (as one would do with a conventional aircraft) nor in the vertical one (as one would do with a pure helicopter).

An example of aircraft of "tail-sitter" type is described in US-2712420-A. The aircraft described in US-2712420-A is equipped with two counter-rotating rotors (to annul the reaction torque to the crankshaft developed by the rotation of the vanes), with variable pitch (to vary the operating conditions of the rotors and then to manage the thrust and the absorbed power). The aircraft is further equipped with a fuselage, wherein the subsystems are housed, and a single wing constrained to the fuselage, finalized to the lift generation in horizontal flight. The equilibrium of the horizontal flight is guaranteed by horizontal planes placed below a yaw damper, directed downwards too. The aircraft is controlled by the pilot by managing the thrust and the deflection of the elevons and of the horizontal small plane, and such logic is applied both in vertical and horizontal flight. The aircraft takes-off in vertical attitude dragged by the thrust of the rotors, it rises as far as a certain quote and subsequently, thanks to the pilot's action on the horizontal small plane and on the elevons, it rotates around the pitching axis by making that the horizontal speed increases so that the wing acquires the capability of generating lift. It is to be noted that such vehicle has rotors with high load on the disc so as to generate the induced speed necessary to make authoritative the aerodynamic controlling surfaces even in horizontal attitude. Such choice, even if dictated by the design specifications, limits considerably the autonomy in hovering of the aircraft due to the very high consumption due to the high loads on the disc. The manoeuvre concludes when the aircraft keeps constant altitude and horizontal attitude with the main wing generating the requested lift and the rotors generating the needed thrust. The control strategy, providing the blowing of the mobile aerodynamic surfaces the aircraft is equipped with, does not allow using rotors with large diameter and then it penalizes the consumption. Moreover the attitude control is strongly penalized by such control strategy: the aerodynamic forces exerted by the wing blown by the rotor are relatively low, they act with a high delay time and do not allow a precise control around all axes of the aircraft.

The object of the present invention is then to overcome the problems illustrated above and this is obtained by means of an aircraft as defined by claim 1. Preferred features of the present invention are subject of the depending claims.

The technical problem placed and solved by the present invention is then to provide a remotely piloted aircraft with better performances having the following advantages: The aircraft according to the present invention joins in a single new aircraft the features of a helicopter (rotors equipped with a collective and cyclic pitch, very low load on the disc) and the features of the aircrafts with not-planar closed wing (the so- called "best wing system", the absolutely most efficient wing system at subsonic speeds).

The aircraft according to the present invention is valid independently from the type of adopted propulsion (the thrust is provided by the propeller, but the power generation can be assigned to propulsors of various nature) and from scale effects: consequently it is applied to all remotely piloted, electrically-propelled, hydrogen, solar and thermal aircrafts, with any size.

The aircraft according to the present invention further allows obtaining the capabilities of vertical take-off and landing, the high efficiency in hovering, typical of a helicopter, the high efficiency in advanced flight, typical of an aircraft with strongly not-planar and closed wing, in a single technological solution. The present invention, in fact, has equal efficiency to that of a helicopter in the flight with fixed point and higher efficiency than that of an aircraft with fixed wing in the horizontal translated flight and during manoeuvre.

Thanks to the control logic of the rotors/propellers under all flight and aircraft conditions, the aircraft according to the present invention is more stable and more manoeuvrable than an aircraft in conventional configuration of equal class, although it is equipped with only two mobile surfaces of aerodynamic control placed on the wing (elevons) and it is still controllable in effective way without any mobile surface of aerodynamic control.

The present invention shows how obtaining a considerable decrease in consumption (increase in autonomy) with respect to a helicopter or a conventional aircraft with fixed wing of the same class, under the respective conditions of nominal operation: in fact, the particular wing architecture of the present invention allows a considerable decrease in resistance induced during the cruising and manoeuvring phase, a considerable reduction in the structural weights (useful in general, but particularly in hovering and in the vertical accelerated flight), as (the sizes and the constructive technology being equal) it is intrinsically more robust and immune from aeroelastic problems in the whole considered flight envelope (with respect to aircrafts/helicopters of the same category and weight class, currently existing on the market).

The speed (mediated on the disc area) of the fluid downwards the rotor disc, induced by the disc itself due to the traction effect, assumes the following formula: wherein T is the thrust (the weight divided by two for a twin-engined aeroplane), A the area of the single rotor disc, p the air density and K a constant equalling 1 at the disc and 2 at the infinite downwards the disc. As it is simple to note, the thrust (or the weight of the aircraft) being equal, the speed induced by the disc increases upon decreasing the disc area: therefore, if one wants to plan efficient means in hovering, indisputably very low speeds will be obtained, insufficient for a precise control in absence of cyclic pitch (the thrust being equal, the higher is the speed induced to the disc, less efficient is the rotor).

The aerodynamic control surface, as the wings themselves, is then "blown" and can exploit such effect to keep the control capability even in the vertical flight, in a more efficient way than the conventional "tail-sitter" of the state of art.

Such advantage derives from the fact of having positioned the wing equipped with control surface more downwards (higher induced speed=greater authority of the control surfaces) than the conventional aircrafts in the vertical take-off currently described in the state of art.

It is important further emphasizing that the conventional "tail-sitter", by exploiting an architecture generally based upon the whole wing, cannot space out too much the wing from the rotor disc due to problems linked to the positioning of the centre of gravity (too ahead engines lead the centre of gravity of the aircraft too ahead than the fourth of the aerodynamic mean chord: in order to compensate such effect one should position the masses downwards the wing with consequent aerodynamic and structural problems).

Other advantages, together with the features and the use modes of the present invention, will result evident from the following detailed description of some preferred embodiments thereof, shown by way of example and not for limitative purpose, by referring to the figures of the enclosed drawings, wherein:

figure 1 is a right rear isometric view of the aircraft according to a first embodiment of the present invention;

figure 2 shows a front view of the aircraft of the embodiment of figure 1 ; figure 3 is a right rear isometric view of the aircraft according to a second embodiment of the present invention;

figure 4 shows a front view of the aircraft of the embodiment of figure 3;

figure 5 shows the aircraft of figures 1-4 in vertical flight (for sake of display simplicity the fitting portions 14 are not represented). In figure the vector weight of the aircraft (V) and the vector thrust of the aircraft (S) are represented;

figure 6 shows the aircraft of figures 1-4 in transition flight (for sake of display simplicity the fitting portions 14 are not represented). In figure the component of vertical thrust (C) and horizontal thrust (O) and the axis of thrust (A-A) and the tilting thereof with respect to the axis of the longitudinal body are represented.

figure 7 shows closed wing with negative stagger (the front wing is the bottom one, the centre of gravity is circled in red and it lies on the plane of the front wing);

figure 8 shows closed wing with positive stagger (the front wing is the top one, the centre of gravity is circled in red and it lies on the plane of the front wing);

figure 9 shows a plan view of the pressure distribution on the two wings with negative and positive stagger (the stagger sign does not distinguish in a plan view);

the graph of figure 10 shows the Lift vs. angle of attack of the three types of compared wings, the continuous curve in dark grey is the one related to the aircraft of the present invention. The wing with positive stagger develops higher lift, the angle of attack and speed being equal, with respect to the other architectures (null stagger and negative staggefy

the graph of figure 11 shows the Resistance induced upon varying the lift coefficient for the three compared wings. The continuous curve in dark grey is the one related to the aircraft of the present invention. The wing with positive stagger develops lower induced resistance, the lift coefficient and speed being equal, with respect to the other architectures (null stagger and negative staggefy

the graph of figure 12 shows the Factors of Oswald or Span-Efficiency Factors for the three wings under examination. The continuous curve in dark grey is the one related to the aircraft of the present invention. The wing with positive stagger has higher efficiency, the angle of attack and speed being equal, with respect to the other architectures (null stagger and negative staggefy

the graph of figure 13 shows the Factors of Oswald or Span-Efficiency Factors for the two wings equipped with stagger divided by the Span-Efficiency Factor of the wing without stagger. The continuous curve is the one related to the present invention. The wing with positive stagger has a higher Span-Efficiency Factor than the one with negative stagger by about 20% and it can be increased by increasing the gap;

the graph of figure 14 shows the course of the moment curves for the three wings under examination. The continuous curve in dark grey is the one related to the aircraft of the present invention. The wing with positive stagger is the only one intrinsically stable among the analysed three ones, the gap being equal. No patent document, nor existing in the specific literature, shows such result;

figure 15 shows the detailed aerodynamic model of the aircraft related to the embodiment represented in figures 16-20;

figure 16 shows a plan view of the aircraft according to a third embodiment of the present invention;

figure 17 shows a side view of the aircraft according to the embodiment of figure 16; figure 18 shows a front view of the aircraft according to the embodiment of figure 16; figure 19 shows a right rear isometric view of the aircraft according to the embodiment of figure 16;

figure 20 shows a right front isometric view of the aircraft according to the embodiment of figure 16.

The present invention will be described hereinafter in details by referring to the above-mentioned figures.

By firstly referring to figure 1 , this represents a perspective view showing as a whole an embodiment of the aircraft according to the present invention. In figure the fuselage 2 is indicated, wherein the direction (L) of longitudinal development and the height (H) between the top profile and the bottom profile of the fuselage itself are highlighted.

As represented in figure, the aircraft according to the present invention comprises two wings 3 with closed structure, wherein each wing 3 in turn comprises a front wing portion 12 substantially with planar shape and a rear wing portion 13 with substantially planar shape too. On each portion of front planar wing 12 a tractive propeller engine 4 is positioned.

The two-engined propeller configuration of the herein described aircraft allows increasing the autonomy by eliminating a monitoring problem around the longitudinal axis, fundamental in vertical flight.

The sizes of the tractive rotors 4 advantageously will be large with respect to the sizes of the aircraft 1 , for example the radius (r) of the rotor could be from 70 cm to 3 m, preferably the radius (r) of the rotor will have sizes approximatively equal to the net halfspan of the wings or the distance of the rotor from the fuselage 2. By considering for the moment aspects related to the vertical flight only, the big size of the two rotors

4 with respect to the whole aircraft 1 makes that the load on the disc (the weight of the aircraft divided by the sum of the surfaces of the disc of the two rotors) is extremely low: this allows maximizing the autonomy in hovenng (the specific power requested to the two propulsors is low) and at the same time the uselessness of additional mechanical devices destined to annul the reaction torque generated by the propellers on the crankshafts, the rotors being counter-rotating according to the direction designated in figure 2. In absence of a double rotor, in fact, the autonomy could be increased only thanks to the increase in the radius of a single rotor, fastened to a propulsor for example on the nose of the fuselage. Such radius increase would lead even to an increase in the reaction torque generated around the axis passing through the centre of the rotor, requesting then the balancing of the torque itself by means of suitable devices known in the state of art.

According to an embodiment the rotors of the aircraft are equipped with mechanisms for varying the collective and cyclic pitch, similar to those used by usual helicopter. Such choice allows the perfect and total controllability of the aircraft in the vertical flight and in the transition one from vertical to horizontal. In fact, by referring to figure 5, it is possible understanding first of all the way in which the control of the longitudinal attitude of the aircraft in vertical position takes place. The mechanism for varying the cyclic pitch allows orienting the axis of thrust by tilting the rotor disc (or the pivoting plate), around anyone of the diameters thereof: this principle makes that in vertical attitude the equilibrium is obtained by tilting the pivoting plate so as to make the axis of thrust to pass by the centre of gravity of the aircraft, whereas when it is necessary to generate monitoring moments, the axis of thrust is made to pass outside the centre of gravity of the aircraft. The same principle is valid for monitoring the attitude around other axes of the aircraft (yawing and rolling). Moreover, the speed monitoring in vertical attitude is performed through the variation of the collective pitch, as in an usual helicopter. In substance: the collective pitch monitors the thrust amount, the cyclic pitch monitors the thrust orientation. The transition to the horizontal flight ideally takes place according to what illustrated in figure 6: the aircraft 1 gathers speed thanks to the tilting of the axis of thrust (or of the pivoting plate/rotor disc) ahead as far as the speed sufficient to make developing the lift at the rear wing which in this way, and by using even the deflection downwards of the monitoring surfaces lifts the aircraft stern by consequently lowering the nose and addressing the thrust vector definitely ahead. When the aircraft is in horizontal support flight, the control for varying the cyclic pitch is not used and the aircraft attitude monitoring takes place only by using the aerodynamic controlling surfaces.

Still by referring to the embodiment of figure 1 , this shows for each wing a fitting portion 14 fitting the front and rear planar wing portions 12, 13 at the terminal ends thereof.

The aircraft according to the present invention further has a first connecting portion 1 1 connecting the front wing 12 to the fuselage 2, and a second connecting portion 15 connecting the rear wing portion 13 to the fuselage 2, preferably the rear wing will be connected to the terminal end of the fuselage 12. According to an embodiment the front wing intersects the fuselage 2, preferably in a median position and still more preferably in central position with respect to the longitudinal axis of the fuselage 2. According to the embodiment of figure 1 the connecting portion 15 is perpendicular to the terminal end of the fuselage 12 and to the rear wing portion 13.

As shown in figures 1-4, the front 12 and rear 13 planar wing portions are staggered both along the direction (L) of longitudinal development of the fuselage and along the height thereof (H), furthermore the rear planar wing portion 13 results in lower position than the planar front wing portion 12 and than the bottom portion of the fuselage 2.

Figures 3 and 4 and 16-20 show a second and a third embodiment wherein the connecting portion 15 of the rear wing 13 to the fuselage 2 is constituted by an element not arranged perpendicularly with respect to the fuselage 2 and to the rear wing portion 13, the tilting of the connecting portion 15 with respect to the rear wing portion 13 preferably will be comprised between 30° and 90°.

According to a preferred embodiment the rear planar wing portion 13 is placed at a distance from the central axis of the fuselage 2 substantially identical to the radius of the rotor or of the semi-wingspan. According to other embodiments this distance could vary from a minimum equal to just over the radius of the fuselage 12 up to a maximum substantially equal the radius of the rotor or of the semi-wingspan, according to other embodiments this distance will be comprised between 0.1 and 1 m.

The connecting portion 15 could have any shape suitable to connect the wing rear portion 13 to the fuselage 12, for example it could be constituted by an element substantially shaped like a zed, sigmoidal or vertical.

The aircraft according to the present invention has different fundamental aspects having a considerable commercial interest thereamong:

• it is capable of taking off/flying/landing vertically;

• it has reduces sizes, the wing surfaces being equal;

• it is more robust, then the robustness being equal, it is lighter;

• it is more efficient according to the gap and the stagger of the second wing with respect to the first one (but still in lower position); • it has high stability with respect to an aircraft both with conventional wing and with closed wing equipped with negative stagger (most studies in progress use negative staggeή.

The aircraft of the present invention could be used in all applications of the remotely piloted aircrafts currently used both in civil and military field: such as for example territory recognition, monitoring of geographic areas, aerial photograph, search and rescue.

The aircraft according to the present invention could be monitored in hovering according to the following logic (described in body axes):

1) pitching: symmetric cyclic.

2) rolling: antisymmetric cyclic

3) yawing: antisymmetric collective

4) rise: symmetric collective

the aircraft transition could be performed by tilting ahead the cyclic controls, the aircraft gathers speed in almost vertical attitude until the second wing starts developing lift, to pass then to the horizontal attitude.

The aircraft in horizontal flight could be monitored according to the following logic (described in body axes):

1) pitching: symmetric control on elevons.

2) rolling: antisymmetric control on elevons.

3) yawing: antisymmetric collective.

Experimental data

Herebelow the numeral results are discussed based upon aerodynamic simulations related to three different closed wings which will be compared to demonstrate the superiority of the configuration of the present invention. The reference closed wing is the one commonly called wing without stagger. To the results of the aerodynamic analysis performed on the wing without stagger the results are added related to two wings with absolute value of stagger equal to two aerodynamic mean chords, the one with negative stagger (front wing in lower position) and the other one with positive stagger (rear wing in lower position, subject of the present invention).

The present invention, with respect to what reported in the publications US5899409 and US20060144991 (wherein one speaks about compromise between stability and efficiency), allows obtaining high efficiency at the same time with high stability without conflicts.

It is possible demonstrating that between the wing with null stagger, the one with negative stagger and the one with positive stagger (all with gap being equal), in terms of aerodynamic efficiency the best solution is the one with positive stagger. The aerodynamic efficiency increases for the wing with positive stagger as, speed and incidence being equal, it develops higher lift (see figure 10) and less induced resistance (see figure 11), thus allowing to obtain very high factors of span-efficiency (see figure 12). If the Oswald factors (or Span-Efficiency Factors) of the two wings with stagger, normalized with respect to that without stagger, are compared, (the gap being equal, in this case gap=0.2), efficiency gains of 20% are obtained (see figure 13).

In substance the above-mentioned closed biplane wing with rear wing passing in lower position than the fuselage, is higher than 20% or more in efficiency, with respect to what reported in the publications US5899409 and US20060144991 , without considering the advantages which can be obtained in terms of stability.

For what concerns the longitudinal static stability, the (opened or closed) biplane wings equipped with considerable gap (higher than 10% of the span) enjoy an intrinsically different property from any other wing architecture: the longitudinal static stability can vary with the incidence as the tilting of the moment curves is not constant.

In case of aircrafts with closed biplane wing equipped with rear wing in upper position

(negative staggeή the stability decreases upon increasing the incidence whereas wings with positive stagger have a positive static stability which increases upon increasing the incidence (see figure 14).

This phenomenon is due to the fact that the rear wing in lower position gives stability upon increasing gradually the incidence of the aircraft, thanks to the fact that the lift and the resistance, projected on the axes X, Z of the aircraft body, tend to develop diving (stabilizing) moments, due to the considerable distance of the aerodynamic centre of the rear wing with respect to the centre of gravity of the aircraft. Such property is advantageous to the purpose of the aircraft flight mechanics: with low incidence the aircraft is very manoeuvrable, whereas with high incidence (low speeds) the aircraft is very stable.

Even if one does not want to link the present invention to any scientific explanation, the main (not single) explanation to the fact that the closed wing with positive stagger is aerodynamically more efficient than that with negative stagger lies in the distribution of the downwash generated by the front wing: generally, for reasonable values of the stagger, the rear wing develops less lift and slightly more resistance if it is in higher position than the front one and the contrary is valid for the wing with positive stagger.

Such effect is also determined by the arrow of the vertical wings: for the positive staggerthe arrow is positive and it contributes to increase the lift on the rear wing.

The present invention has been sofar described by referring to preferred embodiments. It is to be meant that other embodiments belonging to the same inventive core may exist, as defined by the protection scope of the herebelow reported claims.