EP3007974A1

EP3007974A1 - An aircraft having a coanda effect propulsion apparatus

Info

Publication number: EP3007974A1
Application number: EP14744161.2A
Authority: EP
Inventors: Paolo Bellezza Quater; Alfredo Capuani
Original assignee: Nimbus SRL
Current assignee: Nimbus SRL
Priority date: 2013-06-14
Filing date: 2014-06-13
Publication date: 2016-04-20
Also published as: WO2014199351A1; ITTO20130495A1

Abstract

Described herein is an aircraft (1; 1', 100) including : - a fuselage (2; 2'; 102); - a pair of wings (3, 4; 103, 104); - a propulsion apparatus (13; 13'; 113) configured for generating a propulsive thrust by processing a flow of fluid, in particular, air, and configured for varying the direction of said propulsive thrust by the Coanda effect; and - an equipment for generating lift (6, 7; 106, 107; 113, 113') arranged at a nose section (5; 5'; 105) in a substantially end position on said fuselage (2; 2'; 102).

Description

"An aircraft having a Coanda effect propulsion

apparatus"

* * * *

TEXT OF THE DESCRIPTION

Field of the invention

The present invention relates to aircraft that comprise:

- a fuselage;

- a pair of wings; and

- a propulsion apparatus configured for generating a propulsive thrust by processing a flow of fluid, and configured for varying the direction of said propulsive thrust by the Coanda effect.

Aircrafts of this type are known, for example, from EP-A-0078245 and EP-A-0356601.

Prior art

Aircrafts of the known type, for example those forming the subject of the aforementioned documents, include a propulsion apparatus configured for directing the discharged fluid flow over a dorsal surface of a wing profile, in particular, in the proximity of the trailing edge of the wing profile, so as to generate an additional lift (with respect to the lift generated by the wing) thanks to the supercirculation induced on the wing profile itself and to the deflection downwards of the fluid stream by the Coanda effect immediately downstream of the wing profile.

This defines a sort of single-surface ejector, where the primary fluid is constituted by the flow discharged by the propulsion apparatus, the secondary fluid is constituted by the relative motion of the air entrained, by mixing, by the jet along the wing profile, and the subsequent expansion, mixing, and recompression areas are defined by a single active surface that is constituted by the back of the wing profile .

The aforesaid aircrafts are moreover equipped with a mobile surface, which is positioned immediately downstream of the wing profile and can be driven in rotation to obtain a further deflection (by Coanda effect) of the fluid flow discharged by said propulsion apparatus and generate further increase of lift, for example during the landing phase.

However, in such aircrafts problems of stability during flight and of manoeuvreability still remain. In certain cases, maintenance of a stable flying trim imposes the need to forgo ample possibilities of manoeuvre or even proves hard to achieve.

In many aircraft of a known type, in fact, the increase in lift due to supercirculation on the extrados of the wing profile generates a resultant of forces with respect to the centre of gravity of the aircraft that may prove difficult to compensate.

Object of the invention

The object of the invention is to solve the aforementioned technical problems. In particular, the object of the invention is to improve the manoeuvreability of aircraft with Coanda-effect propulsion apparatus of a known type.

Summary of the invention

The object of the invention is achieved by an aircraft having the features forming the subject of one or more of the following claims, which form an integral part of the technical disclosure provided herein in relation to the invention. In particular, the object of the invention is achieved by an aircraft having all the features referred to at the beginning of the present description and further including an equipment for generating lift arranged at a nose section in a substantially end position on the aforesaid fuselage. Brief description of the drawings

The invention will now be described with reference to the attached figures, which are provided purely by way of non-limiting example and wherein:

Figure 1 is a partially sectioned side view of an aircraft according to a first embodiment of the invention;

Figure 2 is a view according to the arrow II of Figure 1 and partially sectioned;

Figure 2A is an enlarged front view according to the arrow IIA of Figure 2;

Figure 2B is a schematic perspective view representing with a thicker line the internal structure of the propulsion apparatus 13, whilst represented by a thinner line is the outline of the structure of the aircraft ;

Figure 3 is a side view of an aircraft according to a second embodiment of the invention;

- Figure 4 is a view according to the arrow IV of Figure 3; and

Figure 5 is a side view of an aircraft according to a third embodiment of the invention.

Detailed description

The reference number 1 in Figure 1 designates as a whole an aircraft according to a first preferred embodiment of the invention. The aircraft 1 is preferably of the unmanned type with remote control.

With reference to Figures 1 and 2, the aircraft 1 includes a fuselage 2 and a wing including a pair of half-wings 3, 4. Provided at the nose 5 of the aircraft is a pair of canard wings 6, 7, whereas at the tail 8 of the aircraft the fuselage forks into two tail beams 9, 10 each bearing a fin 11, 12, basically configured as a vertical tailplane. In any case, in some variants the fins 11, 12 may also be inclined by with respect to the vertical.

The aircraft 1 may moreover be equipped with a landing gear, in particular, of a four-wheel type (necessary to maintain the stability during take-off under the effect of the propulsive thrust), which includes a main gear MG and a nose gear NG .

The aircraft 1 comprises a propulsion apparatus 13, provided within the fuselage 1, including a flow duct 14 having an inlet port I and an outlet port 0.

With reference to Figures 1, 2A, and 2B, set within the flow duct 14 is a shaped diaphragm 15 that defines a bifurcation of the flow duct 14 downstream of the inlet port I. The bifurcation thus defined includes a first duct (first branch) 16 and a second duct (second branch) 17 located on opposite sides of the shaped diaphragm 15.

Arranged, respectively, within the ducts 16, 17 are a first axial-flow motor and a second axial-flow motor with ducted fan driven via a preferably electric motor .

The motors are designated by the reference numbers 18, 19, respectively, and have axes X18, X19 preferably parallel and set on top of one another, both belonging to the longitudinal plane of symmetry of the propulsion apparatus 13 (namely, the plane of the drawing, coinciding also with the longitudinal plane of symmetry of the aircraft 1) . Furthermore, in the preferred embodiment of Figure 1, the motors 18, 19 have a different diameter (in particular, the motor 18 has a smaller diameter than the motor 19) and are located substantially aligned in an axial direction in a position not coinciding with an axis Zl, on which the centre of gravity CG of the aircraft is located.

Variants are possible wherein the axes X18, X19 of the motors 18, 19 once again belong to the longitudinal plane of symmetry of the aircraft 1 but are incident to one another (the motors in any case maintain the stacked arrangement) . Possibly, it is preferable in this configuration to arrange the motors axially staggered to a greater or lesser extent according to the needs. In these variants, the propulsion apparatus 13 is characterized by a greater compactness.

With reference to Figures 1 and 2B, each duct 16, 17 includes two stretches, namely:

a first stretch 16E, 17E (respectively) , housed in which is the corresponding motor 18, 19 and which includes a diverging nozzle that initially has a circular cross section (which reproduces the front shape of the motor 18, 19) and terminates with a rectangular cross section; and

a second, outflow, stretch 16F, 17F, with a rectangular cross section and variable height (i.e., the cross section of each duct 16F, 17F is variable in the longitudinal direction of the duct but is constant in the transverse direction) in so far as it is delimited vertically by an alternation of concave and convex surfaces and laterally, preferentially, by plane walls (in other words, the area of the cross section is variable in the longitudinal direction of the duct); the aforesaid concave and convex surfaces comprise the surfaces of the shaped diaphragm 15 between the two ducts 16, 17 and a first surface 147A and a second surface 14B in the bottom and top parts of the fuselage, respectively.

In this way, three main portions of the flow duct 14, namely (Figures 1, 2B) are defined:

a first, inlet, portion 141, which basically develops from the inlet port I up to the two motors 18, 19; a second portion 14E occupied by the motors 18, 19 and by the respective divergent nozzles 16E, 17E; and

a third, exhaust, portion 14F, which develops substantially in positions corresponding to the outflow stretches 16F, 17F.

The (mobile and fixed) control surfaces of the aircraft 1 will now be described. With reference to Figures 1 and 2, each half-wing 3, 4 includes a respective mobile surface 20, 21 configured for operating as flaperon or elevon (so as to be able to manage the movements of roll of the aircraft 1) . Adjacent to each flaperon or elevon 20, 21, the half- wing 3, 4 has its tip that is bent downwards to improve the lateral-directional stability of the aircraft and may in turn be provided with mobile surfaces.

Each flaperon or elevon 20, 21 can be driven in rotation about a corresponding axis X20, X21 and is movable, with respect to the resting position illustrated in Figure 2 and with a dashed line in Figure 1, by a first angle a_FL as regards the negative- incidence operating range (reduction in lift) , and by a second angle _FL as regards the positive-incidence operating range (increase in lift) of the half-wing.

The canard wings 6, 7 are provided with a respective mobile surface 22, 23, which can be driven in rotation about a respective axis X22, X23 and preferentially can be actuated in a range of angles β₀Α that are preferentially comprised in the range from - 10° to +20° with respect to the resting position so as to modulate the lift generated thereby.

Preferentially, in the resting position (zero incidence of the mobile surfaces 22, 23), the canard wings are in any case oriented so as to have a positive incidence (in particular, approximately 5°). Preferentially, even though it can be envisaged to have embodiments in which both the half-wings 3, 4 and the canard wings 6, 7 are provided with mobile control surfaces, the provision of mobile surfaces on the canard wings 6, 7 has priority. This means that embodiments provided with mobile surfaces on the canard wings 6, 7 and possibly on the half-wings 3, 4 are preferred, whereas, instead, embodiments with mobile surfaces on just the half-wings 3, 4 do not constitute a preferential choice.

In fact, the mobile surfaces 22, 23 on the canard wings are configured for operation as flaps and ailerons, whereas the mobile surfaces on the half-wings are preferably configured for operation as ailerons, which explains the preference in positioning the mobile surfaces on the canard wings at the expense of the half-wings. Furthermore, the priority in installing mobile control surfaces is in favour of the canard since, with mobile control surfaces only on the half- wings 3, 4 and with canard wings 6, 7 having fixed geometry, it would be very difficult to counter the deflected propulsive thrust in landing conditions.

As further mobile control surface, a variable- incidence flap 24 is provided, i.e. which is rotatable around an axis X24 and can be driven in rotation between a first end position and a second end position, represented, respectively, with a solid line and a dashed and double-dotted line in Figure 1. The flap 24 is designed as a complement of the thrust control configuration by Coanda effect to harmonize the needs for surfaces with more marked curvature in the landing phase and for surfaces with a less marked curvature in flight, avoiding situations of high basic resistance.

The first end position corresponds to a cruise flight condition, wherein the extrados of the flap 24 substantially prolongs the surface 14A, without forming any marked deviation with respect thereto (note, in this connection, that the extrados of the flap 24 has an orientation inclined by an angle a_c with respect to the horizontal, equal to 15-20°, like the surface 14A) , whereas the second end position corresponds to a condition of low-speed flight, for example during the landing manoeuvre, where it is necessary to generate further lift. For this reason, in the second end position the extrados of the flap 24 creates a significant deviation with respect to the surface 14A; note, in this connection, that the extrados of the flap 24 has an orientation inclined by an angle _L (greater than ac) with respect to the horizontal, preferentially comprised in the 15°-50° range. The two end positions of the flap 24 are obtained by driving it in rotation in an angular range a₂4 preferentially comprised in the range 0°-50°, preferably 0°-30°.

Moreover provided on board the aircraft 1 is an electronic control unit (not illustrated) , operatively connected to which are the motors 18, 19 and all the mobile control surfaces, as well as. possible sensors on board the aircraft 1. This electronic control unit is pre-arranged for communication with a remote-control device, for example a radiofrequency device, for instance a remote control.

Operation of the aircraft 1 is described in what follows .

The propulsion apparatus 13 is substantially of the type described in the patent application No. RE2011A000049 and identified by the acronym A.C.H.E.O.N. Its operation will not be described in detail as regards the underlying physics, but only the essential aspects thereof will be recalled.

The propulsion apparatus 13 is configured for varying the direction of the air flow processed by the motors 18, 19 by Coanda effect, without mobile parts inside. Each motor 18, 19 draws in a respective air flow, which will thus have a corresponding speed, through the inlet port I .

The air flow traverses the flow duct 14 and is discharged through the outlet port 0 in a direction that, with respect to a horizontal longitudinal axis, may be aligned or slightly inclined upwards or downwards. This depends upon the fact that the stream of fluid exiting from the port O substantially remains on the longitudinal axis of the aircraft 1 or adheres, to a greater or lesser extent, by Coanda effect to the walls 14B or 14A, respectively.

The fact that the stream of fluid exiting from the port 0 adheres to one wall or the other depends upon the speed (and more in general upon the momentum, i.e., upon the mass flow rate) of the fluid in each of the two ducts 16, 17.

In particular, the higher the speed of the fluid

(and more in general the momentum, that is, the mass flow rate) in the ducts 17, 16 respectively, the more markedly the fluid vein adheres to one of the two walls 14A, 14B. In other words, the fluid vein tends to adhere by the Coanda effect to the wall (14A or 14B) belonging to the duct (17 or 16) in which the speed of the fluid (and consequently, the momentum and mass flow rate) is higher.

This moreover means that the adherence of the fluid vein to one wall or the other depends upon the speed of rotation of the two rotors of the motors 18, 19. When the vein adheres to one wall or the other it entrains along with it also the fluid flow that flows in the branch closer to the opposite wall.

By modulating the speed of rotation of the two motors 18, 19 (which are hence actuated independently) it is possible to determine the direction of the flowrate of the discharged fluid.

When the two flowrates in the ducts 16, 17 are identical, the flowrate discharged by the port 0, and consequently the resultant of the thrust exerted by the propulsion apparatus, will be substantially aligned with the longitudinal axis of the aircraft.

As may be noted, in the aircraft 1 the propulsion apparatus 13 is designed in such a way that the thrust exerted thereby will have a direction prevalently facing downwards; note, in this connection the greater extent of the surface 14A at the exhaust port 0 and the larger dimensions of the motor 19. The reason for this is that the flow along the surface 14A and along the flap 24 by the Coanda effect generates a supercirculation on the surface 14A itself (which locally presents a convex conformation resembling the extrados of a wing surface) , which generates an additional lift on the aircraft 1.

In conditions of leveled flight (cruising), the inclination of the thrust of the propulsion apparatus 13 downwards will be at its minimum, with the flap 24 in retracted position (solid line in Figure 1) . In the case where there is the need to land, for example, the thrust of the propulsion apparatus 13 can be deflected more markedly downwards, by increasing the speed of the motor 19 thus getting the fluid flow to adhere markedly to the surface 14A (which in itself increases supercirculation over the surface 14A) and possibly by bringing the flap 24 into its extended position in such a way that, by Coanda effect, the fluid flow is deflected further, up to an angle a_L with respect to the horizontal.

On the opposite side of the fuselage, the canard wings (which, as has been said, preferentially present positive incidence in resting conditions) contribute to stabilizing the flight of the aircraft 1, also generating at the nose a positive lift or a negative lift that can be modulated to maintain the centre of thrust of the aircraft 1 as much as possible corresponding to its centre of gravity CG. In fact, as compared to known aircraft in which generation and modulation of the lift is substantially concentrated on the half-wings behind the centre of gravity, with stabilizing negative lift in the tail area of the aircraft, in the aircraft 1 according to the invention the provision of canard wings with positive incidence makes available two areas of generation of lift at the nose and at the tail of the aircraft.

The manoeuvreability of the aircraft 1 is considerably better than in known aircrafts: not only can the propulsive thrust be vectored (and furthermore without the aid of mobile parts) , but the contributions of lift generated by the canard wings 6, 7, by the wing (i.e., the half-wings 3, 4) and by the propulsion apparatus can be harmonized according to the manoeuvre to be performed. For instance, if the manoeuvre mainly requires the use of the propulsion apparatus, the canard wings and the mobile surfaces of the wing can be used for exerting an action of compensation once again in order to maintain the centre of thrust of the aircraft 1 as much as possible corresponding to the centre of gravity CG.

In practice, this solution affords - as regards the control of the aircraft 1 - a greater number of degrees of freedom as compared to known aircraft.

The mobile surfaces 22, 23 on the canard wings control the pitch movement of the aircraft 1, together with the propulsion apparatus 13, and of course the roll movement. In the case where there are also mobile surfaces 20, 21 on the half-wings 3, 4, these are used preferentially just for control of the roll movement (in combination with, or as an alternative to, the canard wings 6, 7).

With reference to Figures 3 and 4, a second embodiment of an aircraft according to the invention is designated by the reference number 1' . All the components that are identical or practically identical to those already described are designated by the same reference numbers. The components that differ substantially are designated by the same reference numbers already used for the aircraft 1, followed by a prime sign ( ' ) .

The aircraft 1' illustrated in Figure 3 is of the unmanned type with remote control, but variants may also be envisaged wherein the aircraft 1' has a cockpit for a pilot (and possibly a further crew member) . The aircraft 1 includes a fuselage 2' and a wing including the half-wings 3, 4. Provided at the nose 5' of the aircraft are the canard wings 6, 7, whilst at the tail 8' the fuselage forks into the two tail beams 9, 10, each bearing a fin 11, basically configured as a vertical tailplane. In some variants, the fin 11, 12 may also be inclined with respect to the vertical.

The aircraft 1' is moreover equipped with a preferably four-wheel landing gear including the main gear MG and the nose gear NG.

The aircraft 1' includes a propulsion apparatus 13' , which comprises a pod structure 13A' , which carries a motor (an electric motor or a combustion engine) for driving a rotor, in particular a propeller 13B' , in rotation about an axis of rotation X13' . The propeller 13B' may be of the free type (as illustrated in Figure 3) or of the ducted type. Also forming part of the propulsion apparatus is a flow surface 14', which has a convex shape, like the extrados of a wing profile and onto which the propulsion apparatus directs the flow of fluid processed during operation thanks to the orientation of the axis X13.

The (mobile and fixed) control surfaces of the aircraft 1' will now be described. Where they have already been extensively described with reference to the aircraft 1, the control surfaces will only be recalled without repeating the description thereof.

With reference to Figures 3, 4, in a way similar to the aircraft 1, the half-wings 3, 4 include a respective mobile surface 20, 21 configured for operation as flaperon or elevon (so as to be able to manage the movements of roll and pitch of the aircraft 1) . Adjacent to each flaperon or elevon 20, 21, the half-wing 3, 4 has its tip that is bent downwards so as to improve the lateral-directional stability and is also possibly provided with mobile control surfaces.

Each flaperon or elevon 20, 21 can be driven in rotation about a corresponding axis X20, X21 and is movable, with respect to the resting position illustrated in Figure 4, in the way already described for the aircraft 1.

The canard wings 6, 7 are equipped with the mobile control surfaces 22, 23 and preferably have positive incidence ( _CAN is approximately 5°) in resting conditions. The mobile surfaces 22, 23 may be driven in rotation about the respective axes X22, X23 in the same range of angles β₀ΑΝ described previously so as to modulate the lift generated by the canard wings 6, 7.

Preferentially, even though it can be envisaged to have embodiments wherein both the half-wings 3, 4 and the canard wings 6, 7, are provided with mobile control surfaces, provision of mobile surfaces on the canard wings 6, 7 has priority. This means that the embodiments provided with mobile surfaces on the canard wings 6, 7 and possibly on the half-wings 3, 4 are preferred, whereas embodiments with mobile surfaces on just the half-wings 3, 4 do not constitute a preferential choice. The reasons have already been set forth above in relation to the aircraft 1.

As further mobile control surface, a variable- incidence flap 24' is provided, i.e., a flap that can turn about an axis X24' and can be driven in rotation between a first end position and a second end position, which are indicated by a solid line and a dashed line, respectively, in Figure 3.

The first end position corresponds to a cruise flight condition, where the extrados of the flap 24' substantially prolongs the surface 14' , without forming any marked deviation with respect thereto (note, in this connection, that the extrados of the flap 24' has an orientation inclined by an angle a_c' _t with respect to the horizontal, that is approximately 10°-15°), as in the case of the surface 14', whereas the second end position corresponds to a condition of low-speed flight, for example during a landing manoeuvre, where it is necessary to generate further lift. For this reason, in the second end position the extrados of the flap 24' creates a significant deviation with respect to the surface 14' (note, in this connection, that the extrados of the flap 24' has an orientation inclined by an angle α_τ > a_c' with respect to the horizontal, that is approximately 30°-40°) .

Moreover provided on board the aircraft 1' is an electronic control unit (not illustrated) , operatively connected to which are the propulsion apparatus 13' and all the mobile control surfaces, as well as possible sensors on board the aircraft 1' . As in the case of the aircraft 1, this electronic control unit can be prearranged for communication with a remote-control device, for example a radiofrequency remote-control device, for instance, a remote control.

Operation of the aircraft 1' is described in what follows .

The propulsion apparatus 13' generates a propulsive thrust by driving in rotation the rotor 13B' , which processes a flow of fluid, which is directed onto the surface 14' so as to induce a supercirculation of the fluid thereover, which results in the generation of an additional lift. Via the flap 24', the fluid vein that flows away along the surface 14' is then deflected by the Coanda effect so as to further modulate the additional flow generated via the propulsion apparatus 13' .

As in the case of the flap 24 on the aircraft 1, the flap 24' can be driven in rotation about the axis X24' between a first operative end position and a second operative end position. In the first operative position, indicated by a solid line in Figure 3, the flap 24' is substantially aligned with the surface 14', this meaning a position such that no significant deviation of the stream of fluid is obtained with respect to the direction determined by the flow over the surface 14', this position moreover being associated to a condition of cruise flight.

Instead, in the second position - indicated by a dashed line in Figure 3 - the flap 24' is characterized by a marked inclination with respect to the surface 14', so as to form a significant deviation by the Coanda effect of the stream of fluid that flows over the surface 14', generating a maximum additional lift, which is useful, for example, during a low-speed manoeuvre, such as the landing manoeuvre. The position of the flap 24' can be modulated between the two end positions described above, thereby obtaining a variation of the direction of the fluid flow (and hence of the propulsive thrust) discharged by the propulsion apparatus 13' thanks to the Coanda effect, and a consequent modulation of the lift.

According to an advantageous aspect of the invention, a feature that characterizes the aircraft 1' is the integration of the propulsive flows drawn in by the propeller 13B' with the circulation generated by the wing, which results in an increase in the overall lift due to the superposition of four effects:

i) basic circulation of the wing due to incidence; ii) deflection of the propulsive thrust of the propeller 13B' due to the Coanda effect, with creation of a vertical component of lift;

iii) increase in the negative pressure on the back as a result of intake of flow by the propeller 13B' ; and

iv) increase in circulation of the wing due to the curtain effect of the deflected central flow, which substantially acts as a flap.

Note that, unlike the use of jet engines (turbojets or turbofans), which typically generate a thrust that is largely concentrated and localized in space, the choice of the propeller 13B' enables generation of a flow with higher rate over the entire central part of the wing profile, thus increasing the section affected by the flow itself and the spatial distribution of the thrust. If a traditional jet engine (turbojet or turbofan) were to be used, this would impose the use of the single flow of the engine, moreover with the need to set the engine itself in a very advanced position along the fuselage so as to entrain the surrounding air and favour mixing thereof with the propulsive flow before this detaches from the fuselage. Instead, the propeller 13B' achieves the above result, even though it is set in a set-back position on the fuselage, thanks to its intake capacity, which generates a propulsive flow of larger section right from the start.

According to a further advantageous aspect of the invention, the shape of the wing of the aircraft 1' in plan view is important, with the central area affected by the action of the propeller 13B' set in an advanced position with respect to the trailing edge of the wing itself. This makes it possible to bring the resultant of the deflected propulsive thrust closer to the centre of gravity CG' , avoiding pitch moments that may be difficult to control (and balance) using the canard wings 6, 7 alone.

As regards manoeuvreability and stability of the aircraft 1' during manoeuvring, all the considerations expressed regarding the aircraft 1 apply. In particular, the canard wings 6, 7 contribute to improving the quality of flight and the manoeuvreability of the aircraft 1' in the way already described in regard to the aircraft 1; i.e., they contribute to stabilizing the flight of the aircraft 1' also generating at the nose a positive lift or a negative lift that can be modulated to maintain the centre of thrust of the aircraft 1' as much as possible at its centre of gravity CG.

The mobile surfaces 22, 23 on the canard wings control the pitch movement of the aircraft 1 together with the propulsion apparatus 13' , and of course the roll movement. In the case where there are also the mobile surfaces 20, 21 on the half-wings 3, 4, these are used preferentially just for control of the roll movement (in combination with, or as an alternative to, the canard wings 6, 7) .

With reference to Figure 5, a third embodiment of an aircraft according to the invention is designated by the reference number 100 in Figure 5. The reference numbers in brackets are associated to an element (or a geometrical reference) that is identical to, and symmetrical with respect to, the one illustrated in the figure, but is located on the opposite side of the aircraft (and is consequently not visible) .

The aircraft 100 illustrated in Figure 5 is of the unmanned type with remote control, but variants may also be envisaged wherein the aircraft 1' has a cockpit for a pilot (and possibly a further crew member) . The aircraft 100 includes a fuselage 102 and a wing comprising a pair of half-wings 103, 104. At the nose 105, the aircraft is provided with a pair of canard wings 106, 107, whereas at the tail 108 of the aircraft the fuselage forks into two tail beams 109,110 each bearing a fin 111, 112, basically configured as a vertical tailplane. In any case, in some variants, the fins 111, 112 may also be inclined with respect to the vertical .

The aircraft 100 is moreover equipped with a preferentially four-wheel landing, including a main gear MG and a nose gear NG.

The aircraft 1 comprises a propulsion apparatus 113 including two motors 113' installed on the canard wings 106. 107. Each motor 113' is basically identical save for the reduced scale - to the propulsion apparatus 13. In other words, each motor 113 constitutes a version on a smaller scale of the propulsion apparatus 13 of the aircraft 1.

In the case in point, each motor 113' includes a flow duct 114 having an inlet port I and an outlet port 0. Set within the flow duct 114 is a shaped diaphragm 115, which defines a bifurcation of the flow duct 114 downstream of the inlet port I . The branching thus defined includes a first duct (first branch) 116 and a second duct (second branch) 117, which are located on opposite sides of the shaped diaphragm 115.

Arranged within the ducts 116, 117, respectively, are a first axial-flow motor and a second axial-flow motor with ducted fan driven by a preferably electric motor.

The motors are designated by the reference numbers 118, 119 respectively, and have axes X118, X119 preferably parallel to and set on top of one another, both belonging to the longitudinal plane of symmetry of the motor 113' . Furthermore, in the preferred embodiment of Figure 5 the motors 118, 119 have a different diameter (in particular, the motor 118 has a smaller diameter than the motor 119) and are located substantially aligned in an axial direction in a position not coinciding with an axis Z100 on which the centre of gravity CG of the aircraft is located.

Moreover variants are possible wherein the axes X118, X119 of the motors 118, 119 once again belong to the longitudinal plane of symmetry of the motor 113', but are incident (the motors in any case maintain the arrangement where they are set on top of one another) . Possibly, it is preferable in this configuration to arrange the motors axially staggered, to a greater or lesser extent according to the needs. In these variants, the motor 113' is characterized by a greater compactness .

Each duct 116, 117 includes two stretches (see, for further reference, the previous description, given the identity of the two systems, save for a scale factor) , namely: a first stretch 116E, 117E (respectively) , housed in which is the corresponding motor 118, 119 and which includes a diverging nozzle that initially has a circular cross section (which reproduces the front shape of the motor 118, 119) and terminates with a rectangular cross section; and

a second, outflow, stretch 116F, 117F, with a rectangular cross section and variable height (i.e., the cross section of each duct 116F, 117F is variable in the longitudinal direction of the duct but is constant in the transverse direction) in so far as it is delimited vertically by an alternation of concave and convex surfaces and laterally, preferentially, by plane walls (in other words, the area of the section is variable in the longitudinal direction of the duct); these concave and convex surfaces comprise the surfaces of the shaped diaphragm 115 that separates the two ducts 116, 117 and a first and a second surface 114A, 114B in the bottom and top parts, respectively, of the fuselage.

Three main portions of the flow duct 114 are in this way defined, namely:

a first, inlet, portion 1141, which develops basically from the inlet port I up to the two motors 118, 119;

a second portion 114E occupied by the motors 118, 119 and by the respective divergent nozzles 116E, 117E; and

a third, discharge, portion 114F, which develops substantially in a region corresponding to the outflow stretches 116F, 117F.

The (mobile and fixed) control surfaces of the aircraft 100 will now be described. With reference to Figure 5, each half-wing 103, 104 includes a respective mobile surface 120, 121 configured for operating as flaperon or elevon (so as to be able to manage the movements of roll of the aircraft 1) . Adjacent to each flaperon or elevon 120, 121, the half-wing 103, 104 has its tip that is bent downwards to improve the lateral- directional stability of the aircraft and is possibly equipped with a further mobile control surface.

Each flaperon or elevon 120, 121 can be driven in rotation about a corresponding axis X120, X121, as has already been described for the flaperon 20, 21, also as regards the angular ranges provided therefor.

The canard wings 106, 107 are preferably oriented in order to have a positive incidence CX_CAN in resting conditions (so as to balance the aircraft with a further amount of lift at the nose) comprised in the angular range 0°-20°.

The canard wings 106, 107 are equipped with a respective mobile surface 122, 123 that can be driven in rotation about a respective axis X122, X123 and preferentially that can be actuated in a range of angles identical to the one already been indicated for the aircraft 1.

Preferentially, even though it is possible to envisage embodiments in which both the half-wings 103, 104 and the canard wings 106, 107 are provided with mobile control surfaces, the provision of mobile surfaces on the canard wings 106, 107 has priority. This means that the embodiments provided with mobile surfaces on the canard wings 106, 107 and possibly on the half-wings 103, 104 are preferred, whereas embodiments with mobile surfaces on just the half-wings 103, 104 do not constitute a preferential choice.

Note, however, that it is possible to envisage some variants in which the aircraft 100 is equipped, instead of with the canard wings 106, 107, with just the motors 113'. In this case, the canard wings, which on all the aircraft described herein constitute an equipment for generation of lift at the nose, as well as for the modulation thereof (in so far as they are provided with mobile control surfaces), wherein this equipment is set at the nose in a substantially end position on the fuselage, can be functionally replaced by modulation of the angle of thrust of the motors 113, obtained by electronic control of the motors 118, 119.

Finally, provided on board the aircraft 100 is an electronic control unit (not illustrated) , to which the motors 118, 119 and all the mobile control surfaces are operatively connected, as well as possible sensors on board the aircraft 100. This electronic control unit can be arranged for communication with a remote-control device, for example, a radiofrequency remote-control device, for instance, a remote control.

Operation of the aircraft 100 is described in what follows .

Each of the two motors 113' presents an operation altogether identical to that of the propulsion apparatus 13 already described. For this reason, only the essential features of operation of the latter will be recalled.

Each motor 113' is configured for varying the direction of the flow of air processed by the motors 118, 119 by the Coanda effect, without mobile parts inside. Each motor 118, 119 draws in a respective air flow, which will thus have a corresponding speed, through the inlet port I. The air flow traverses the flow duct 114 and is discharged through the outlet port 0 with a variable direction, according to whether the stream of fluid exiting from the port 0 remains substantially along the longitudinal axis of the motor 113 or adheres to a greater or lesser extent, by the Coanda effect, to the walls 114B or 114A, respectively. This depends, as has already been seen, upon the speed (and more in general upon the momentum, i.e., upon the mass flow rate) of the fluid in each of the two branches 116, 117, and hence, in the final analysis, upon the speed of rotation of the rotors of the motors 118, 119. When the vein adheres to one wall or the other, it entrains along with it also the fluid flow that flows in the branch that is closer to the opposite wall .

By modulating the speed of rotation of the two motors 118, 119 (which are hence driven independently), it is possible to determine the direction of the flow of outgoing fluid thus obtaining a corresponding modulation of the direction of the propulsive thrust. Illustrated in Figure 5 are three directions representing as many operating ranges of the motors 113' :

a first configuration of thrust (cruise thrust) envisages that the resultant of the propulsive force generated by each motor 113 has an inclination Yi, with respect to a horizontal longitudinal axis of the aircraft 100, of between 0° and 5°;

a second configuration of thrust (take-off thrust) envisages that the resultant of the propulsive force generated by each motor 113 has an inclination Y2, with respect to a horizontal longitudinal axis of the aircraft 100, of between 10° and 15°; and

a third configuration of thrust (landing thrust or loitering thrust) envisages that the resultant of the propulsive force generated by each motor 113' has an inclination γ₃, with respect to a horizontal longitudinal axis of the aircraft 100, of between 45° and 50°.

Finally, as regards the manoeuvreability of the aircraft 100 and the advantages deriving from the provision of the canard wings, substantially all the considerations set forth above with regard to the aircraft 1 and 1' apply. Nonetheless, the fact of locating the propulsion apparatus in the nose region of the aircraft contributes even more to the distribution of the forces (of propulsion and lift) along the fuselage of the aircraft, contributing to accentuating all the benefits that have already been discussed.

Of course, the details of construction and the embodiments may vary widely with respect to what has been described and illustrated herein, without thereby departing from the scope of protection of the present invention, as defined by the annexed claims.

Claims

1. An aircraft (1; 1' ; 100) including:

- a fuselage (2; 2'; 102),

- a pair of wings (4, 3; 103, 104),

- a propulsion system (13; 13'; 113) configured for generating a thrust and configured for the variation of the direction of said thrust by Coanda effect, and

- an equipment for generating lift (6, 7; 106, 107; 113, 113') arranged at a nose section (5; 5'; 105) in a substantially end position on said fuselage (2; 2' ; 102) .

2. The aircraft (1; 1'; 100) according to Claim 1, wherein said equipment for generating lift (6, 7; 106, 107; 113, 113') is further configured for the modulation of the lift generated thereby.

3. The aircraft (1; 1'; 100) according to any of claims 1, 2, wherein said equipment for generating lift (6, 7; 106, 107; 113, 113') includes a pair of canard wings (6, 7; 106, 107), each canard wing (6, 7; 106, 107) of said pair including a mobile control surface (22, 23; 122, 123) .

4. The aircraft (100) according to any of the previous claims, wherein said equipment for generating lift (106, 107; 113, 113') includes a first and a second motor (113', 113') constituting said propulsion system ( 113 ) .

5. The aircraft (1; 100) according to any of the previous claims, wherein said propulsion system includes:

- a flow duct (14, 114) having an inlet port (I) and an outlet port (0) ,

a shaped septum (15, 115) which provides a bifurcation of the flow duct (14, 114), and

- a first and a second axial-flow motors (18, 19; 5 118, 119) arranged in a correspondent duct (16, 17;

116, 117) constituting a respective branch of said bifurcation .

6. The aircraft according to Claim 5, wherein said second axial-flow motor (19; 119) has a larger diameter

10 than said first axial-flow motor (18; 118) .

7. The aircraft (1; 100) according to Claim 5 or Claim 6, wherein each duct (16, 17; 116, 117) includes:

- a first stretch (16E, 17E; 116E, 117E) wherein the corresponding axial-flow motor (18, 19; 118, 119)

15. is arranged and including a divergent nozzle terminating with a rectangular cross section,

- a second outflow stretch (16F, 17F, 116F, 117F) , having a cross section with rectangular shape and variable area.

20 8. The aircraft (1) according to any of Claims 5 to 7, wherein said propulsion system is provided within said fuselage (2) of the aircraft (1).

9. The aircraft according to any of Claims 4 to 7, wherein said first and second motors (113) are each

25 located on a corresponding canard wing (106, 17) of said pair.

10. The aircraft (1; 1') according to Claim 1, wherein said propulsion system (13; 13') is configured for directing, during operation, a fluid flow processed

30 thereby towards a surface (14A, 14') shaped in a convex fashion so as to generate an additional lift by supercirculation over said surface shaped in a convex fashion (14A, 14'), said aircraft (1; 1') further including a flap (24; 24') arranged in correspondence of a trailing edge of said surface (14A; 14') and having variable incidence to modify the thrust direction by Coanda effect.