“CONVERTIBLE AIRCRAFT CAPABLE OF HOVERING” Cross-Reference to Related Applications This Patent Application claims priority from European Patent Application No. 22174103.6 filed on May 18, 2022 the entire disclosure of which is incorporated herein by reference. Technical Field The present invention relates to an aircraft that is convertible between a first configuration of hovering flight or of flight along a predominantly vertical trajectory and a second forward flight or cruise configuration. Background In the aviation sector, aeroplanes are normally used for high cruising speeds, in particular greater than 150 knots and high altitudes, e.g. above 30,000 feet. At cruising speeds and high altitudes, aeroplanes use fixed wings to generate the lift necessary to sustain the plane. A sufficient value of this lift can only be obtained by accelerating the aeroplane on runways of considerable length. These runways are also necessary to allow the same aeroplanes to land. In contrast, helicopters normally have lower cruising speeds than aeroplanes and generate the necessary lift for sustenance through the rotation of the main rotor blades. As a result, helicopters can land/take off without the need for a horizontal speed and using particularly small surfaces. Moreover, helicopters are able to hover and to fly at relatively low altitudes and speeds, resulting thus as particularly manoeuvrable and suitable for demanding manoeuvres such as rescuing people in the mountains or at sea. Nevertheless, helicopters have inherent limitations in terms of maximum operational altitude, which is around 20000 feet, and maximum operational speed, which cannot exceed 150 knots. In order to meet the demand for aircrafts that have the same manoeuvrability and flexibility of use as the helicopter and at the same time overcome the inherent limitations mentioned above, convertiplanes are known and which constitute a type of convertible aircraft. An example of a convertiplane is described in patent application US-B-10,011,349. In more detail, the convertiplane described in the aforesaid application essentially comprises: - a fuselage extending along a first longitudinal axis; and - a pair of half-wings projecting cantilevered from respective parts of the fuselage that are opposite to each other, and having respective free ends opposite to the fuselage and aligned along a second transverse axis that is substantially orthogonal to the first longitudinal axis. The convertiplane further comprises: - a pair of nacelles housing the respective motors; and - a pair of rotors which are rotatable around respective third axes and operatively connected to respective motors. The rotors are inclinable with respect to the wing around a fourth axis, preferably parallel to the second axis. The convertiplanes are also able to selectively assume: - a first “helicopter” configuration, wherein the rotors are arranged with the respective third axes that are substantially vertical and orthogonal to the first axis of the convertiplane and orthogonal to the respective motors; or - a second “aeroplane” configuration, wherein the rotors are arranged with respective third axes that are substantially parallel to the first axis of the same convertiplane and coaxial to the respective engines. The convertiplane further comprises a tail portion provided with a pair of aerodynamic surfaces adapted to generate a second adjustable lift/downforce value that ensures a desired degree of longitudinal stability to the convertiplane in the second configuration. Recently, several proposals for convertible aircrafts with electric propulsion have been developed, for example of the type described in patent application WO-A-2020/105045. US-A-2021/253234 discloses a convertible aircraft according to the preamble of claim 1. US-A-2020/277045 discloses a vertical take-off and landing (VTOL) aircraft having three lifting surfaces and separate lift and cruise systems. The VTOL aircraft may include a fuselage having a roll axis, a thrust rotor to produce a propulsion thrust, first and second rotor booms, first and second canard surfaces, first and second wing surfaces, first and second tail surfaces, and a plurality of lift rotors to produce a lifting thrust force. The plurality of lift rotors includes a first plurality of lift rotors positioned on the first rotor boom and a second plurality of lift rotors positioned on the second rotor boom. The first and second rotor booms may be substantially parallel to the roll axis of the fuselage, where the fuselage is positioned between the first and second rotor booms. Each of the first and second rotor booms may be secured to the aircraft at three locations. US2020/108918 discloses an aircraft motor includes a bearing assembly including a first plurality of rotor alignment magnets; a magnet support structure fixedly mounted on a shaft of the motor in a spaced apart relation to the bearing assembly, the magnet support structure including a second plurality of rotor alignment magnets such that when the vertical thrust engine is disengaged, attraction between the first and second rotor alignment magnets causes the magnet support structure to rotate relative to the bearing assembly to an alignment position defined by the relative placement of north and south poles of the first and second plurality of rotor alignment magnets. US-B-9120560 discloses a vertical take-off and landing aircraft includes a fixed wing airframe having opposed left and right wings extending from left and right sides, respectively, of a fuselage having opposed leading and trailing extremities and an empennage located behind the trailing extremity. Four fixed, open and horizontal, vertical take-off and landing (VTOL) thrust rotors are mounted to the airframe in a quadrotor pattern for providing vertical lift to the aircraft, and a vertical, forward thrust rotor is mounted to the trailing extremity of the fuselage between the trailing extremity of the fuselage and the empennage for providing forward thrust to the aircraft. The four VTOL thrust rotors are coplanar being and operating in a common plane that is parallel relative to, and being level with, top surfaces of the left and right wings in and around a region of each of the four VTOL thrust rotors. US-A-2021/197965 discloses an aircraft with a fuselage module and at least two vertical lift rotor modules supporting at least four rotor assemblies. Each rotor assembly is supported by a rotor boom having at least one boom free end and a boom mounting portion. Each rotor assembly has at least one vertical lift rotor mounted on the boom free end. Each boom mounting portion is removably couplable to the fuselage module. The vertical lift rotor modules are configured such that when coupled to the fuselage module, a pair of the rotor assemblies are located on each of laterally opposite sides of the fuselage module, and the rotor assemblies of each pair are respectively located forward of and aft of a wing center portion. A pair of wings are configured to be removably couplable to the wing center portion. The aircraft includes a forward thrust module removably couplable to the fuselage body. There is a perceived need in the sector to realize a convertible aircraft for sports competition or for use in a personal air mobility environment, which exceeds the limits normally set by the certification requirements for aircraft intended for passenger transport. In particular, there is a perceived need in the sector to ensure a high degree of stability of the aircraft while reducing the overall aerodynamic drag. Summary Aim of the present invention is to realize a convertible aircraft which allows to satisfy at least one of the needs specified above in a simple and economical way. According to the invention, this aim is achieved by a convertible aircraft as claimed in claim 1. Brief Description of the Drawings For a better understanding of the present invention, three preferred non-limiting embodiments are described, by way of example only and with the aid of the accompanying drawings, wherein: - Figure 1 is a perspective view from above and according to a first angle of view of a convertible aircraft, realized according to a first embodiment of the present invention; - Figure 2 shows the aircraft of Figure 1 in a top view; - Figure 3 shows the aircraft of Figure 1 in a front view; - Figure 4 shows the fuselage of the aircraft of Figure 1 in a side view, with parts removed for clarity’s sake; - Figures 5 and 6 are cross-sections along the lines V-V and VI-VI of the fuselage of Figure 4; - Figure 7 is a perspective view of a fuselage of the aircraft of Figure 1; - Figure 8 is a perspective view of the aircraft of Figures 1 to 7, with first components in transparency for clarity’s sake; - Figure 9 is a perspective view of the aircraft of Figures 1 to 7, with second components in transparency for clarity’s sake; - Figures 10 and 11 are front views of the aircraft of Figures 1 to 9, respectively with third and fourth components not shown for clarity’s sake; - Figure 12 is a functional diagram of some components of the aircraft of Figures 1 to 11; - Figure 13 is a side view of a fuselage of an aircraft realized according to a second embodiment of the invention, with parts removed for clarity’s sake; and - Figure 14 is a cross-section of a fuselage of an aircraft realized according to a third embodiment of the invention, with parts removed for clarity’s sake. Description of Embodiments With reference to Figures 1 to 11, 1 denotes an aircraft capable of hovering. In greater detail, the aircraft 1 is selectively switchable between: - a first configuration (Figures 1, 2 and 3), wherein it performs a hovering manoeuvre or moves forward along a predominantly vertical trajectory; and - a second configuration (Figure 8 and 9), wherein it is in a forward flight condition and proceeds along a mainly horizontal trajectory. It must be specified that in the following present disclosure, expressions such as “above”, “below”, “at the front”, “behind” and the like are used to refer to a normal forward or “hovering” flight condition of the convertiplane 1 shown in Figures 1, 2, 3, 8 and 9. It is possible to identify a set of axes integral to the aircraft 1 and originating in a centre of gravity of the aircraft 1 itself formed by: - a longitudinal axis Y of the aircraft 1 itself; - an axis X orthogonal to the axis Y; and an axis Z orthogonal to the axes X, Y. In a known manner, the rotations of the aircraft 1 around the axes Y, X, Z are associated with the following manoeuvres: - roll, i.e. rotation around the axis Y; - pitching, i.e. rotation around the axis X; - yaw, i.e. rotation around the axis Z. More particularly, the aircraft 1 is intended for sports competitions or for use in a personal air mobility environment. The aircraft 1 essentially comprises: - a fuselage 2 which is elongated along the axis Y and defining a nose 4 and a tail 5 of the aircraft 1; - a pair of half-wings 3 extending cantilevered from respective mutually opposite sidewalls 19 of the fuselage 2 and transversely to the axis Y; and - a tail portion 6 projecting cantilevered from the tail 5 of the fuselage 2 transversely to the fuselage 2. With reference to the normal forward flight operating conditions, the aircraft 1 proceeds in a direction oriented from the tail 5 to the nose 4 parallel to the axis Y. The half-wings 3 are intended to provide a first lift value to the aircraft 1 which is adapted to sustain the aircraft 1 arranged in the second configuration. The half-wings 3 comprise respective free ends 15 opposite the fuselage 2. The half-wings 3 extend above the fuselage 2. The half-wings 3 are divergent from each other proceeding from the fuselage 2 towards the respective free ends 15. In the case shown, the half-wings 3 comprise: - respective root portions 11 projecting cantilevered from respective sidewalls 19; and - respective end portions 12 arranged above the fuselage 2 and defining the respective ends 15 of the half-wings 3 themselves. The tail portion 6 comprises, in turn, an aerodynamic surface 8 adapted to generate a second lift/downforce value to ensure a desired degree of longitudinal stability to the aircraft 1 itself arranged in the second configuration. Preferably, the half-wings 3 have negative deflection and/or negative dihedral angle. The half-wings 3 are, moreover, without movable aerodynamic surfaces, such as appendages or flaps. Preferably, the aircraft 1 further comprises a pair of canard-type aerodynamic surfaces 9 projecting cantilevered from respective mutually opposite sides of the nose 4 of the fuselage 2 and adapted to generate a third lift/downforce value to ensure the desired degree of longitudinal stability to the aircraft 1 itself arranged in the second configuration. The aerodynamic surfaces 9 comprise, in turn: - respective root portions 17 projecting cantilevered from and connected to respective sidewalls 19 of the fuselage 2; and - respective end portions 18 arranged on the opposite side of the corresponding root portions 17 with respect to the fuselage 2. In particular, the root 17 and end 18 portions are inclined to each other (Figure 3). In the case shown, the wingspan of the half-wings 3 is greater than the wingspan of the aerodynamic surface 8. The wingspan of the aerodynamic surfaces 9 is greater than the wingspan of the aerodynamic surface 8. In this description, the term “wingspan” means the distance between opposite free ends 15, of the respective half-wings 3 and opposite free ends of the aerodynamic surfaces 8, 9. The aerodynamic surfaces 9 are arranged below the half-wings 3 (Figure 3). The fuselage 2 further comprises (Figures 4 to 7): - a pair of portions 100, 101 arranged one after the other, proceeding along the axis Y from the nose 4 towards the tail 5; and - an electric power source 102 (only schematically shown in Figures 4, 6 and 8) arranged inside the portion 101 of the fuselage 2. The source 102 could comprise a plurality of electric batteries, a plurality of fuel cells, or an electric generator. The portion 100 defines a housing 104 for a first crew member 105. The portions 100, 101 respectively define a first and a second section 110, 111 (Figures 5, 6 and 11) in a plane orthogonal to the axis Y. The portion 101 further comprises a pair of air intakes 120 that are open towards the outside of the fuselage 2, towards the nose 4, and fluidically connected with the source 102 so as to convey, in use, a flow of cooling air onto the source 102 itself, following the forward motion of aircraft 1. The section 111 has a larger area than the section 110. The air intakes 120 are arranged externally to said portion 100 and on respective mutually opposite sidewalls 19 of the portion 100, in a view of the aircraft 1 parallel to the axis Y. The fuselage 2 houses a conveying duct 109 (Figure 4) adapted to convey fresh cooling air onto the source 102 and fluidically connected with the air intakes 120. The section 111 is symmetrical with respect to the axes Z, Y and comprises, in turn (Figure 6): - an area 130; and - an area 131 with greater extension than the area 130, housing the source 102 and defining said air intakes 120. The area 131 is arranged below the area 130 with respect to the axis Z, with reference to the normal forward position of the aircraft 1. The area 131 projects laterally from both sides of the area 130, with reference to the axis X. The fuselage 2 further comprises a transparent dome 140 adapted to cover the housing 104 for the first crew member 105. The dome 140 is carried by the portion 100. The dome 140 can slide or be opened laterally or towards the tail 5, in the case shown. The section 110 is also symmetrical with respect to the axes Y, Z and comprises, in turn (Figure 5): - an area 145 arranged at the area 130 along the axis Y; and - an area 146 arranged at the area 131 along the axis Y and arranged below the area 145 with respect to the axis Z, with reference to the normal forward position of the aircraft 1. The area 145 is contained in the area 130 and the area 146 is contained in the area 131, in a view of the aircraft 1 according to a direction parallel to the axis Y. The section 110 further comprises an area 150 interposed between the areas 145, 146 along the axis Z. The area 145 is shaped to accommodate the head of the first crew member 105 and the area 146 is shaped to support the legs of the aforesaid first crew member 105. The area 150 projects from both sides of the areas 145, 146, with reference to the axis X, so as to define two lateral fairings 142 adapted to accommodate respective arms of the first crew member 105. In this way, the section 100 defines a safety cell for the first crew member 105. The aircraft 1 further comprises (Figures 1): - a portion 107 interposed between the nose 4 and the portion 100 along the axis Y; and - a portion 108 interposed between the portion 101 and the tail 5 along the axis Y. The half-wings 3 are fixed to the portion 108. The portions 107, 108 are delimited above by respective stretches 112, 113. The stretch 112 extends between the domes 140 and the nose 4 along the axis Y. The stretch 112 is inclined with respect to the axis Y and extends below the dome 140 at progressively increasing distances from the dome 140 proceeding towards the nose 4. The stretch 113 extends between the portion 101 and the tail 5 along the axis Y. The stretch 113 is inclined with respect to the axis Y and extends below the portion 101 at progressively increasing distances from the dome 140 proceeding towards the tail 5. The portions 100, 101 are joined together parallel to the axis Y. The fuselage 2, the half-wings 3 and the aerodynamic surfaces 8, 9 have respective foam inserts (not shown). With reference to Figure 2, the aerodynamic surfaces 9 comprise respective foam inserts 201, 202 carried respectively by the corresponding root 17 and end 18 portions. The half-wings 3 comprise respective foam inserts 203, 204 carried respectively by the corresponding root 11 and end 12 portions. The aerodynamic surfaces 8 comprise respective foam inserts 205. The inserts 201, 202; 203, 204; 205 confer structural strength to the aerodynamic surface 9, to the half-wings 3 and to the aerodynamic surface 8, and confer buoyancy capacity to the aircraft 1. The aircraft 1 further comprises (Figures 1 to 3): - a pair of rotors 20a, 20b which are rotatable around respective fixed axes B, C with respect to the fuselage 2; - a pair of rotors 21a, 21b which are rotatable around respective fixed axes D, E with respect to the fuselage 2; and - a pair of rotors 22a, 22b which are rotatable around respective axes F, G and inclinable with respect to an axis H (indicated only in Figure 2) between a first position that is assumed when the aircraft 1 is in the first configuration (Figures 1 to 3) and a second position that is assumed when the aircraft 1 is in the second configuration (Figures 8 and 9). In the shown case, the axes F, G of the rotors 22a, 22b are inclinable with respect to the axis H by more or less fifteen degrees towards the nose 4 or the tail 5 with respect to the axis Z. The aircraft 1 further comprises: - a plurality of electric powertrains 40a, 40b, 40c, 40d, 40e, 40f (only schematically shown in Figure 12) adapted to be electrically powered wholly or partly by the source 102 and to drag into rotation respective rotors 20a, 20b, 21a, 21b, 22a, 22b; and - a control unit 71 (only schematically shown in Figure 12) receiving as input a plurality of control signals provided by the crew or by an autopilot or a remote control system, and programmed to provide as output a plurality of commands to command the electric powertrains 40a, 40b, 40c, 40d, 40e, 40f so that the respective rotors 20a, 20b; 21a, 21b; 22a, 22b provide desired values of the relative thrusts. In greater detail, the control unit 71 is programmed to command the rotors 20a, 20b; 21a, 21b; 22a, 22b to generate respective mutually independent thrusts. More particularly, the control unit 71 is programmed to command the rotors 20a, 20b; 21a, 21b; to generate respective thrusts having a resultant which is parallel to the axis Z, both when the aircraft 1 is in the first configuration and when the aircraft 1 is in the second configuration. The control unit 71 is programmed to command the rotors 20a, 20b; 21a, 21b to generate respective zero thrusts under predetermined operating conditions and when the aircraft 1 is in the first configuration. The axes B, C; D, E and F, G are arranged symmetrically to the axis Y. In the case shown, the axes B, C, D and E are parallel to each other and parallel to the axis Z. The axes B, D, F; C, E, G are aligned with each other orthogonally to the axis Y when the aircraft 1 is arranged in the first configuration. The axes F, G are arranged parallel to the axis Z when the rotors 22a, 22b are arranged in the first position. The axes F, G are arranged orthogonally to the axes B, C; D, E and parallel to the axis Y when the rotors 22a, 22b are arranged in the second position. In one embodiment, the rotors 20a, 20b; 21a, 21b; 22a, 22b are with fixed pitch. The aircraft 1 further comprises (Figures 2): - a pair of supports 30a, 30b in order to support respective rotors 20a, 20b in a fixed manner with respect to respective aerodynamic surfaces 9; - a pair of supports 31a, 31b in order to support respective rotors 21a, 21b in a fixed manner with respect to respective half-wings 3 and to the tail portion 6; and - a pair of supports 32a, 32b in order to support respective rotors 22a, 22b to respective half-wings 3 in an inclinable manner with respect to the axis H. Preferably, the supports 32a, 32b are spaced apart from the ends 15 of the respective half-wings 3, with reference to an extension direction of the same half- wings 3. The supports 30a, 30b are spaced apart from the free ends of the respective aerodynamic surfaces 9, with reference to an extension direction of the same aerodynamic surfaces 9. In particular, the supports 30a, 30b are conformed as respective plates projecting cantilevered below from respective aerodynamic surfaces 9 in front of the nose 4 and lying on respective planes orthogonal to the direction Y and arranged vertically in the normal position of the aircraft 1 (Figure 1). The axes B, C are arranged in front of the nose 4. The axes D, E are arranged in front of the respective aerodynamic surface 8. The rotors 22a, 22b are interposed between the respective half-wings 3 and the nose 4 along the axis Y both when the aircraft 1 is arranged in the first configuration and when the aircraft 1 is arranged in the second configuration. The rotors 22a, 22b are arranged above the respective half-wings 3 when the aircraft 1 is arranged in the first configuration (Figures 1 to 3), and in front of the respective half-wings 3 when the aircraft 1 is arranged in the second configuration (Figures 8 and 9). With particular reference to Figure 9, the tail portion 6 essentially comprises - the aerodynamic surface 8 extending from respective mutually opposite sides of the tail 5 and lying on a plane substantially orthogonal to the axis Z; and Advantageously, the tail portion 6 further comprises a pair of fins 52 arranged transversely to the aerodynamic surface 8; each fin 52 comprises, in turn: - a portion 54 extending from a first side of said aerodynamic surface 8; and - a second portion 55 extending from a second side, opposite the first side of the aerodynamic surface 8; the supports 31a, 31b of each rotor 21a, 21b are furthermore connected to a corresponding said half-wing 3 and to a corresponding fin 52. In more detail, the fins 52 ensure steering stability to the aircraft 1 in the second configuration. The fins 52 project cantilevered from respective opposite ends 53 of the surface 8. The portion 54 extends upwards starting from the relative end 53 with reference to the normal position of the aircraft 1 and lying on a plane inclined with respect to the axes Y, Z; and The portion 55 extends downwards starting from the relative end 53 with reference to the normal position of the aircraft 1, inclined with respect to the portion 54, and lies on a plane orthogonal to the axis X. The supports 31a, 31b are conformed as rods parallel to the axis Y, arranged on respective mutual opposite sides of the fuselage 2, and extending integrally, each, between a respective half-wing 3 and the portion 55 of a respective fin 52. The fins 52 define extensions of the respective supports 31a, 31b parallel to the axis Y. The portions 54, 55 are arranged below the rotors 21a, 21b, with reference to the normal position of the aircraft 1. In particular, the aerodynamic surface 8 comprises: - a main portion 60; and - an appendage 61 (shown only in Figure 8) movably connected to the main portion 60. In the shown case, the appendage 61 is arranged behind the main portion 60 with reference to the normal position of the aircraft 1, and is hinged to the main portion 60 around an axis parallel to the axes H, X. The aircraft 1 further comprises an actuator 65 operable to rotate the appendage 61 with respect to the main portion 60 and around the axis parallel to the axes H, X. The actuator 65 essentially comprises (Figure 8): - a fixed body 66 housed in the fuselage 2 and housing a servomotor 64; - a rod 67 elongated parallel to the axes H, X and fixed to the appendage 61; and - a rod 68 slidable with respect to the fixed body 66 parallel to the direction Y following actuation of the servomotor 64 and hinged, on the opposite side of the fixed body 66, to the rod 67 around the axis parallel to the axes H, X. The aircraft 1 further comprises (Figures 1 and 8): - a pair of devices 45 – for example, skids or carriages carried by respective aerodynamic surfaces 9; and - a pair of devices 46 – for example, skids or carriages carried by respective portions 55 from the fin 52 on the side opposite the surface 8. The devices 45, 46 are flexibly mounted with respect to the aerodynamic surfaces 9 and to the portions 55, respectively. The following describes the operation of the aircraft 1, with reference to a use in a sports competition and/or in a personal/urban air vehicle (PAV) environment. The first crew member 105 accesses the portion 100 through the sliding or laterally openable dome 140. Next, the first crew member 105 is accommodated in the portion 100 of the fuselage 2. The source 102 powers, in whole or in part the powertrains 40a, 40b, 40c, 40d, 40e, 40f with electrical power. The air intakes 120 convey fresh air into the duct 109, which continuously cools the source 102. The control unit 71 receives as input a plurality of control signals provided by the crew, by an autopilot or a remote control system, and provides as output a plurality of commands to command the electric powertrains 40a, 40b, 40c, 40d, 40e, 40f independently of each other so that the respective rotors 20a, 20b; 21a, 21b; 22a, 22b provide desired values of the relative thrusts. The aircraft 1 lands and takes off arranged in the first configuration with the rotors 22a, 22b arranged in the first position wherein the relative thrusts are directed parallel to the axis Z (Figures 8 and 9). The aircraft 1 moves forward cruising in the second configuration with the rotors 22a, 22b arranged in the second position wherein the respective thrusts are arranged parallel to the axis Y (Figures 1, 2 and 3). In the first configuration, the lift required to sustain the aircraft 1 is provided by the rotors 20a, 20b; 21a, 21b and 22a, 22b. In the second configuration, the lift required to sustain the aircraft 1 is mostly provided by the half- wings 3. The rotors 20a, 20b; 21a, 21b can be deactivated, if necessary. The control and the manoeuvrability of the aircraft around the axes X, Y, Z is obtained by appropriately acting on the values of the thrusts generated by the rotors 20a, 20b; 21a, 21b; 22a, 22b and/or on the relative inclination of the rotors 22a, 22b with respect to the axis H. With reference to the second configuration, the variation of the angle of inclination of the appendage 61 through the actuator 65 contributes to controlling and manoeuvring the aircraft 1 with respect to the axis X. The fins 52 ensure steering stability to the aircraft 1 in the second configuration With reference to Figure 13, 2’ denotes a fuselage of an aircraft realized according to a second embodiment of the present invention. The fuselage 2’ is similar to the fuselage 2 and will be described below only in so far as it differs from the latter; equal or equivalent parts of the fuselages 2, 2’ will be marked, where possible, with the same reference numbers. In particular, the fuselage 2’ differs from the fuselage 2 in that the portion 101 defines a housing 104’ for a second crew member 106 arranged behind the housing 104 along the axis Y. In particular, the housing 104’ allows the second member 106 to sit astride at least a front portion of the source 102. The dome 140’ also extends beyond the portion 101 towards the tail 5 so as to be arranged above the housing 104’. The operation of the aircraft 1 with the fuselage 2’ is completely similar to that of the fuselage 2 and is therefore not described in detail. With reference to Figures 14, 2’’ denotes a fuselage of an aircraft realized according to a third embodiment of the present invention. The fuselage 2’’ is similar to the fuselage 2 and will be described below only in so far as it differs from the latter; equal or equivalent parts of the fuselages 2, 2’’ will be marked, where possible, with the same reference numbers. In particular, the fuselage 2’’ differs from the fuselage 2 in that the portion 101 defines a housing 104’’ for a second and a third crew member 106a’’, 106b’’ side by side along the axis X. The section 111’’ is conformed similarly to the section 111 of the fuselage 2 but has increased overall dimensions in order to accommodate the two crew members. In a nutshell, the section 111’’ comprises: - an area 145’’; and - an area 146’’ arranged below the area 146’’, with reference to the normal forward position of the aircraft 1; and - an area 150’’ interposed between the areas 145, 146 along the axis Z. The area 145’’ is shaped to accommodate the heads of the first and second crew members 106a’’, 106b’’ and the area 146’’ is shaped to support the legs of the aforesaid first and second crew member 105, 106. The area 150’’ projects from both sides of the areas 145’’, 146’’, with reference to the axis X, so as to define two lateral fairings 142’’ adapted to accommodate the arms of the first member 105 and of the second member 106 of the crew. The operation of the aircraft 1 with the fuselage 2’’ is completely similar to that of the fuselage 2 and is therefore not described in detail. From an examination of the characteristics of the aircraft 1 according to the present invention, the advantages that it allows obtaining are evident. In greater detail, the fins 52 of the tail portion 6 are arranged transversely to the aerodynamic surface 8, comprise respective portions 54, 55 extending above and below the aerodynamic surface 8 and are connected to the supports 31a, 31b of the rotors 21a, 21b. Therefore, the fins 52 in addition to ensuring lateral-directional stability to the aircraft 1 effectively contribute to mechanically supporting the rotors 21a, 21b. It is thus possible to make the supports 31a, 31b slimmer and lighter, thus reducing the overall aerodynamic drag of the aircraft 1 and consequently improving its sporting performance. The fuselage 2, 2’, 2’’ also houses the fixed body 66 of the actuator 65 adapted to control the position of the appendage 61 with respect to the main portion 60 of the tail plane 50. Consequently, the geometry of the tail portion 6 can be defined without the need to house the aforesaid actuator 65, helping to further reduce the aerodynamic drag of the aircraft 1 compared to the solutions of the known type. Each fin 52 has development almost completely orthogonal to the axis Z and is furthermore arranged below the rotors 21a, 21b, reducing the interference with the downward blow generated by the corresponding rotor 21a, 21b, and consequently improving the behaviour of the aircraft 1 in the first configuration. The fins 52 carry respective carriages 46. It is thus possible to use the same component both to achieve the desired level of lateral-directional stability of the aircraft 1 and to support the carriages 46, without requiring dedicated support elements of the carriages 46 themselves. On the one hand, this further reduces the aerodynamic drag of the aircraft 1. On the other hand, it allows to guarantee the aircraft 1 four support points, two of which are defined by the carriages 45 and two of which are defined by the carriages 46. In this way, the aircraft 1 shows high stability and safety during the manoeuvres of loading/unloading the source 102 and/or of ascent/descent of the crew. It is clear that the aircraft 1 described and shown herein may be subject to modifications and variations without thereby departing from the scope of protection defined by the claims. In particular, the fuselage 2, 2’, 2’’ could comprise a single air intake 120 arranged at a relative sidewall 19. The housing 104 could accommodate two first crew members 105 arranged side by side parallel to the axis X. The aircraft 1 could comprise instead of the rotors 22a, 22b, one or more reaction or jet engines fuelled by fossil fuel, which are fixed with respect to the fuselage 2, and configured to generate a thrust parallel to the axis Y under forward flight conditions. The aircraft 1 may not comprise the aerodynamic surface 8. The axes B, C; D, E may not be parallel to the axis Z and may be inclined with respect to the axis Z by an angle ranging between - 15 and + 15 degrees. In particular, the axes B, C (D, E) could converge in the axis Z above or below the fuselage 2. Some or all of the rotors 20a, 20b, 21a, 21b, 22a, 22b may have variable pitch.