Patent Publication Number: US-2023159181-A1

Title: Aircraft capable of hovering

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This patent application claims priority from Italian patent application no. 20171458.1 filed on 27 Apr. 2020, the entire disclosure of which is incorporated herein by reference. 
     TECHNICAL FIELD 
     The present invention relates to an aircraft capable of hovering, in particular a helicopter or a convertiplane. 
     BACKGROUND ART 
     Helicopters are known essentially comprising a fuselage, a main rotor rotatable around a first axis and arranged on a top portion of the fuselage, and an anti-torque rotor arranged at a tail end of the helicopter and rotatable around a second axis, transverse to the first axis. 
     The helicopters of the known type also comprise a motor system and a transmission group for transmitting the motion from an outlet shaft of the motor system to the main rotor. 
     In greater detail, the motor system, known in the sector as the “turbo-shaft”, creates an open thermodynamic cycle. 
     The motor system comprises:
         a support body;   an air intake adapted to allow the entry of a first flow rate of air;   a compressor fluidically connected to the air intake, supplied with the aforesaid first flow rate of air and adapted to compress said first flow rate of air;   a combustion chamber, in which the first flow rate of compressed air from the compressor is mixed with a second flow rate of fuel and undergoes a combustion process generating a third flow rate of high temperature exhaust gases; and   one or more turbines, inside which the third flow rate of high temperature exhaust gases leaving the combustion chamber expands by driving the compressor and the outlet shaft of the motor system in rotation independently of each other.       

     Each motor system also comprises:
         a converging nozzle arranged downstream of the respective turbine and adapted to accelerate the third flow rate of exhaust gases; and   an exhaust gas discharge duct ending in a respective opening of the support body and inside which the converging nozzle is arranged.       

     The helicopters of the known type also comprise a lubrication system adapted to allow the lubrication of the motor system and to contribute to the cooling of the motor system itself. 
     In more detail, the lubrication system comprises:
         a collection tank for a lubricating fluid, for example oil; and   a distribution circuit configured so as to distribute the lubricating fluid in certain regions of the motor system and to allow the return of said lubricating fluid to the tank.       

     During said circulation, the lubricating fluid comes into contact with the moving components of the motor system and consequently increases its temperature. 
     In order to prevent the temperature of the lubricating fluid from becoming excessive, the lubrication system comprises, in the known solutions:
         a fan adapted to generate a fourth stream of fresh air; and   a heat exchanger that allows to cool the lubricating fluid through the heat exchange with the fourth stream of air generated by the fan.       

     The use of the fan or similar active systems leads to an increase in weight of the helicopter. 
     The actuation of said fan or similar systems of the active type also requires a rate of power, for example electrical, directly extracted from the on-board system, or mechanical, made available by the motor system and which is subtracted from the actual value of mechanical power available to the motor shaft. 
     The fan and the relative driving group also require respective housings inside the helicopter, which sometimes give rise to problems of integration with other systems and/or with the structure and the shape of the fuselage itself. 
     Finally, the fan and the relative driving group are inevitably subject to risks of failure, thus worsening the overall reliability of the helicopter&#39;s motor system. Such solutions of the known type also require more frequent inspection and maintenance intervals, thus increasing the overall operating costs of the helicopter. 
     WO-A-2003/037715 describes a passive cooling system for an auxiliary power unit of an aircraft. 
     The auxiliary power unit is designed to supply the various systems on board the aircraft with electric power and compressed air. 
     The auxiliary power unit is basically a gas turbine system equipped with a pair of compressors housed inside a nacelle. 
     Said auxiliary power unit also comprises a heat exchanger for cooling a lubricating fluid. 
     The gas turbine comprises, in a known manner, a further discharge duct for the mixture of exhaust gases housed inside the nacelle. 
     In a first embodiment, the nacelle defines:
         a single air intake to supply the compressors and the heat exchanger with a stream of air; and   an outlet mouth of the further discharge duct of the gas turbine.       

     The nacelle also houses:
         a first duct arranged downstream of the single air intake; and   a second and a third duct, which define respective branches of the first duct.       

     In greater detail, the second duct extends between the first duct and a first suction mouth of one of the compressors. 
     The third duct branches, in turn, into a first and a second section arranged downstream of the first duct. 
     The first section fluidically connects the first duct and a second suction mouth of the other one of the compressors. 
     The second section fluidically connects the first duct and is arranged inside the internal volume of the nacelle. The heat exchanger is interposed along the second section. 
     The further discharge duct of the gas turbine has a converging nozzle with progressively decreasing areas. The nozzle has, in turn, a downstream section fluidically connected with the internal volume of the nacelle downstream of the second section. 
     In this way, the exhaust gas flow in the discharge duct causes a reduction in the speeds of the exhaust gases and a consequent depressurization at the downstream section of the nozzle and, consequently, within the nacelle regions located downstream of the second section. 
     Said depressurization causes a first stream of air through the heat exchanger and a second stream of air towards the first and second suction mouths of the compressors. 
     In other words, the nozzle defines an ejector at its downstream section. 
     Since the second section is open inside the volume of the nacelle, the first stream of air cools both the heat exchanger and the internal volume of the nacelle. 
     Consequently, it is not possible to independently control the cooling of the exchanger and the cooling of the internal volume of the nacelle. 
     WO-A-2003/037715 describes further embodiments, in each of which the stream of cooling air of the internal volume of the nacelle is driven by the ejector into the discharge duct of the motor system. 
     Consequently, the same stream of air cools both the heat exchanger and the internal volume of the nacelle. 
     A need is felt in the sector to independently and optimally control the cooling of the casing of the motor system and the heat exchanger. 
     This with the aim of optimising the cooling of the nacelle of the motor system and of the heat exchanger from a dimensional and functional point of view. 
     DISCLOSURE OF INVENTION 
     The object of the present invention is the realization of an aircraft capable of hovering, which allows to satisfy the aforesaid requirement in a simple and economic way. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       For a better understanding of the present invention, a preferred embodiment is described below, by way of non-limiting example and with reference to the accompanying drawings, wherein: 
         FIG.  1    illustrates a perspective view of a helicopter according to the invention; 
         FIG.  2    is a view from above and on a strongly enlarged scale of the helicopter of  FIG.  1    and of a relative motor system, on a strongly enlarged scale and with parts removed for clarity&#39;s sake; 
         FIG.  3    is a front perspective view according to a first visual angle and on a further enlarged scale of some components of the motor system of the helicopter of  FIGS.  1  and  2   ; 
         FIG.  4    is an exploded view of the components of the motor system of  FIG.  3   ; 
         FIG.  5    is a rear perspective view according to a further visual angle from  FIG.  3    and on a further enlarged scale of the components of the motor system of  FIGS.  3  and  4   ; 
         FIGS.  6  and  7    illustrate in a perspective view and on a further enlarged scale the components of  FIGS.  3  to  5   , with parts removed for clarity&#39;s sake; 
         FIG.  8    is a rear view of the components of the motor system of  FIGS.  3  to  7   ; 
         FIG.  9    illustrates an exploded view of the components of the motor system of Figures from 1 to 8; 
         FIGS.  10  and  11    respectively illustrate a top and perspective view of the components of the motor system of  FIGS.  3  to  9   ; 
         FIG.  12    is a front view of further components of the motor system of Figures from 1 to 11; 
         FIG.  13    illustrates a further embodiment of the further components of  FIG.  12   ; and 
         FIG.  14    illustrates a further embodiment of the further components of  FIG.  13   . 
     
    
    
     BEST MODE FOR CARRYING OUT THE INVENTION 
     With reference to  FIG.  1   , numeral  1  indicates a helicopter essentially comprising a fuselage  2  provided with a front nose  5 , a main rotor  3  placed at a top of the fuselage  2  and rotatable around a first axis, and a tail rotor  4  carried by a drift protruding from the fuselage  2  on the side opposite the nose  5  and rotatable around a second axis transverse to the first axis. 
     It should be noted that in the following of the present description, expressions such as “above”, “below”, “front”, “rear” and the like are used with reference to advanced flight or “hovering” conditions of the helicopter  1  illustrated in  FIG.  1    and wherein the main rotor  3  is arranged above the fuselage  2  and the nose  5  is arranged in front of the tail rotor  4 . 
     The helicopter  1  comprises a motor system  6  housed in a motor bay  8  delimited by a support body  7 . 
     The motor bay  8  is fluidically connected with an air intake  10  adapted to allow the entry of a stream of cooling air into the motor bay  8  itself. 
     The helicopter  1  also comprises a transmission group (not illustrated as known per se and not part of the present invention) adapted to connect an outlet shaft (also not illustrated) of the motor system  6  to a shaft for driving the main rotor  3  rotatable about an axis A. 
     The motor system  6  behaves like a gas turbine plant realising an open Joule-Brayton thermodynamic cycle. 
     The motor system  6  essentially comprises ( FIG.  2   ):
         an air intake  9  obtained in a flank of the fuselage  2  and through which a first flow rate of low temperature air is sucked into the motor bay  8 ;   a compressor  11  (only schematically illustrated) provided with a suction duct (not illustrated) of the first flow rate of air fluidically connected with the air intake  9  and adapted to compress said first flow rate of air;   a combustion chamber  13  (only schematically illustrated), in which the first flow rate of air compressed by the compressor  11  reacts with a second fuel flow rate generating a third flow rate of high temperature exhaust gases; and   a pair of turbines  14  (only schematically illustrated), inside which the third flow rate of high temperature exhaust gases leaving the combustion chamber  13  expands by driving the compressor  11  and the outlet shaft in rotation.       

     In particular, the compressor  11 , the turbines  14  and the outlet shaft are rotatable around the axis A. 
     The air intake  9  is arranged laterally to the axis A and is distinct from the air intake  10 . 
     The motor system  6  further comprises a duct  17  for discharging the third flow rate of exhaust gases ending in a respective mouth  18  of the support body  7 . 
     The helicopter  1  also comprises a lubrication system (known per se and not illustrated in detail) adapted to allow the lubrication and to contribute to the cooling of the motor system  6 . 
     In greater detail, the lubrication system comprises a collection tank (not illustrated) of a lubricating fluid, a distribution circuit (also not illustrated) configured to distribute the lubricating fluid in certain regions of the motor system  6  and to allow the return of said lubricating fluid into the tank. 
     During said circulation, the lubricating fluid comes into contact with the moving components of the motor system  6  and increases its temperature. 
     The lubrication system further comprises a heat exchanger  20 , which allows to cool the lubricating fluid by means of the heat exchange with a stream of air. 
     In other words, the heat exchanger  20  is a radiator crossed by the lubricating fluid and cooled by the stream of air. 
     The heat exchanger  20  is arranged outside the motor system  6 . 
     The helicopter  1  further comprises:
         a further air intake  25  open on a flank of the fuselage  2  and adapted to suck a fourth flow rate of air; and   a duct  26  along which the heat exchanger  20  is interposed and through which the fourth flow rate of air flows.       

     The air intake  25  is distinct from the air intake  10 . 
     The duct  26 , in turn, comprises:
         an inlet section  27  extending between the air intake  25  and the heat exchanger  20 ; and   a pair of channels  28  ( FIGS.  2 ,  4  and  7  to  10   ) extending from the heat exchanger  20  and arranged on the opposite side of the section  27  with respect to the heat exchanger  20 .       

     The helicopter  1  furthermore comprises a converging nozzle  15  arranged downstream of the turbines  14  and crossed by the third flow rate of exhaust gases. 
     The nozzle  15  has a tubular shape of axis A and comprises:
         a surface  31  radially internal to the axis A and shaped like a tapered cone, running from the turbines  14  towards the discharge duct  17 ; and   a plurality of lobes  32  angularly equally spaced around the axis A and protruding in a cantilever fashion from the surface  31  towards the axis A itself.       

     The nozzle  15  comprises ( FIGS.  3 ,  4  and  6  to  9   ):
         an upstream section  38  fluidically connected with the turbines  14 ; and   a downstream section  39  opposite the upstream section  38  and arranged on the side of the discharge duct  17 .       

     The downstream section  39  of the nozzle  15  is fluidically connected with the channels  28  of the duct  26  and with the discharge duct  17 . 
     With reference to  FIG.  2   , the helicopter  1  comprises an ejector  80  formed by the downstream section  39  of the nozzle  15  and by the channels  28  of the duct  26 . 
     The term ejector or jet-pump means in the present description a pump formed by a converging nozzle inside which a primary stream of a fluid is conveyed and having a downstream section fluidically connected to a duct. The converging shape of the nozzle causes a lowering of the static pressure in the downstream section of the nozzle, which allows to suck a secondary stream through the duct. Said primary and secondary streams mix in the outlet section of the nozzle. 
     The channels  28  comprise respective openings  43  ( FIG.  8   ) opposite the heat exchanger  20  and open at the downstream section  39  of the nozzle  15 . 
     More precisely, the nozzle  15  causes a lowering of the static pressure of the third flow rate of exhaust gases leaving the turbines  14  at the downstream section  39 . Said lowering of static pressure draws the fourth flow rate of air through the duct  26  which cools the heat exchanger  20  and mixes in the downstream section with the third flow rate of exhaust gases giving rise to a fifth flow rate of exhaust gases and air through the downstream section  39 . 
     The nozzle  15  is housed partly inside the nozzle  16  and is arranged upstream of the discharge duct  17 . 
     The downstream section  39  of the nozzle  15  is fluidically connected with the channels  28 . 
     Advantageously, the helicopter  1  comprises a further converging nozzle  16  ( FIGS.  2  to  7  and  9  to  11   ) and having a downstream section  49  fluidically connected with the discharge duct  17  and the motor bay  8 , so as to create a sixth flow rate of cooling air of the motor bay  8  itself directed from the air intake  10  towards the discharge duct  17  and by-passing the motor system  6 . 
     This sixth flow rate of air cools the motor bay  8 . 
     In greater detail, the nozzle  16  comprises an upstream section  48  opposite the downstream section  49  and fluidically connected with the downstream section  39  of the nozzle  15 . 
     The downstream section  49  is fluidically connected with the motor bay  8 , as will be described in more detail below. 
     The helicopter  1  comprises a further ejector  90  formed by the nozzle  16  and the motor bay  8 . 
     More precisely, the nozzle  16  causes a lowering of the static pressure of the fifth flow rate of air and exhaust gases at the downstream section  49 . Said lowering of static pressure draws a sixth flow rate of air through the motor bay  8  which cools the motor bay  8  itself and mixes in the downstream section  49  with the fifth flow rate of exhaust gases and air giving rise to a seventh flow rate of exhaust gases and air through the downstream section  49 . 
     With reference to  FIGS.  3  to  11   , the nozzle  15  is arranged upstream of the nozzle  16 , running from the turbine  14  towards the discharge duct  17 . 
     The nozzles  15 ,  16  are arranged coaxially to the axis A. 
     The nozzle  15  is housed partly inside the nozzle  16  and is arranged upstream of the discharge duct  17 . 
     The nozzle  15  is also radially spaced from the nozzle  16 . 
     With reference to  FIGS.  3  to  8 ,  10  and  11   , the helicopter  1  also comprises a support body  40 , which supports the heat exchanger  20 . 
     The support body  40  integrally defines the channels  28  and the nozzle  16  and houses the nozzle  15 . 
     In greater detail, the support body  40  integrally comprises, running from the turbine  14  towards the discharge duct  17 :
         a portion  44  protruding with respect to the axis A and inside which the nozzle  15  is housed and defining the channels  28  of the duct  16 ; and   a tubular portion  45  with respect to the axis A, housed partly inside the portion  44  and partly inside the discharge duct  17 , and defining therewith the nozzle  16 .       

     The portion  44  supports the heat exchanger  20  and defines the channels  28 . 
     In particular, the portion  44  comprises:
         a discoidal wall  46 , orthogonal to the axis A and delimiting the support body  40  on the side of the turbine  14 ; and   a curved wall  47  protruding in a cantilever fashion from an end edge  70  of the wall  46  radially opposite to the axis A towards the discharge duct  17 .       

     The walls  44 ,  47  surround the portion  45  below and are open above the portion  45 . 
     The wall  46  further comprises an end edge  72  radially internal and opposite the end edge  70 . The nozzle  15  is fixed circumferentially to the end edge  72  ( FIG.  7   ). 
     The wall  46  furthermore comprises an upper end  41  which is rectilinear and orthogonal with respect to the axis A, and is closed below the portion  45 . 
     The wall  47  comprises a pair of upper ends  42 , parallel to the axis A and connected to the end  41 . The wall  47  is also closed below the portion  45 . 
     The support body  40  further comprises ( FIG.  7   ):
         a pair of walls  50  facing respective portions  51  ( FIG.  8   ) of corresponding walls  47 ; and   a wall  52  axially opposed to the wall  46 .       

     The walls  47 ,  50  extend in an axially interposed position between the walls  46 ,  52 . 
     The walls  47 ,  50  extend symmetrically to each other with respect to an axis B orthogonal to the axis A and arranged, in use, vertically. 
     More precisely, each wall  50  comprises:
         an end  57  fixed to the heat exchanger  20 ;   an end  59  that is free and opposite to the respective end  57 .       

     Each wall  52  in turn comprises an end  58  fixed to the heat exchanger  20  and connected to the respective end  57 . 
     The walls  50  define a diverging cusp running from the respective common ends  57  towards the respective ends  59  that are free and spaced apart between them ( FIG.  8   ). 
     The support body  40  defines ( FIG.  4   ) a perimetric end edge  60 , open and fixed to the heat exchanger  20 . 
     The edge  60  is delimited by the end  41  of the wall  46  and by the ends  58  of the walls  57  by respective parts that are axially opposite each other. 
     The edge  60  is also delimited by the ends  42  of the wall  47 . 
     The edge  60  is, in the case illustrated, rectangular. 
     The ends  57  of the walls  50  are arranged parallel to the ends  42  and cross the edge  60 . 
     More particularly, the ends  57  of the walls  50  divide the edge  60  into two equal areas defining respective inlet sections of respective channels  28  opposite the respective openings  43 . 
     The ends  41 ,  58  are axially opposed to each other. 
     The ends  42 ,  57  are opposed to each other and axially interposed between the ends  41 ,  42 . 
     With particular reference to  FIG.  8   , each opening  43  is axially delimited by the end  59  of the respective wall  50  and by an end of the respective wall  52  opposite the end  58 . 
     The channels  28  have a progressively decreasing thickness in an orthogonal direction to the respective walls  50 , running from the ends  57  towards the respective ends  59 , i.e. from the heat exchanger  20  towards the respective openings  43 . 
     The portion  45  comprises, running from the turbines  14  towards the discharge duct  17  ( FIGS.  6 ,  8  and  11   ):
         a wall  65  surrounding the nozzle  15 ; and   a wall  66  axially offset with respect to the nozzle  15 , and provided with a plurality of lobes  67  protruding in a cantilever fashion towards the axis A from a radially internal surface  69  of the wall  66  itself and a plurality of through openings  68 .       

     With particular reference to  FIG.  8   , the wall  65  is shaped like an arc coaxial to the axis A. 
     The wall  65  extends between the ends  59  of the respective walls  51 . 
     The wall  65  surrounds an arcuate section of corresponding angular width of the nozzle  15 . 
     The wall  65  extends, in the case illustrated, over an arc of about ninety degrees and extends symmetrically to an axis B orthogonal to the axis A and arranged vertically in a normal flight configuration of the helicopter  1 . 
     The wall  66  is filleted to wall  52 . 
     The lobes  67  are angularly equally spaced around the axis A and are arranged at the respective lobes  32  of the nozzle  15 , running parallel to the axis A. 
     The openings  68  are angularly equally spaced around the axis A and elongated along the axis A. 
     Each opening  68  is associated with a respective lobe  67 . 
     The lobes  67  protrude in a cantilever fashion from the wall  66  at respective openings  68 . 
     The wall  66  is partially housed inside the discharge duct  17 . 
     More in particular, the discharge duct  17  comprises an annular end  71  opposite the mouth  18 . The end  71  defines an annular groove  81  with the wall  65  axially opposite to the wall  46 . 
     In particular, the discharge duct  17  has, running from the end  71  towards the mouth  18 , a section  73  converging with respect to the axis A, a section  74  with constant diameter and a section  75  diverging with respect to the axis A. 
     The groove  81  and the openings  68  fluidically connect the motor bay  8  with the downstream section  49  of the nozzle  16 . 
     The wall  66  comprises an annular end  82  axially opposite to the wall  46 , housed inside the discharge duct  17  and radially spaced from said discharge duct  17 . 
     In particular ( FIG.  12   ), the portion  44  extends symmetrically to the axis B orthogonal to the axis A and is arranged above the axis A, with reference to the normal operating configuration of the helicopter  1 . The air intake  25  is arranged below the main rotor  3 . In this way, the downward flow generated by the main rotor  3  causes a dynamic stream inside the air intake  25 , which further contributes to cooling the heat exchanger  20 . 
     According to an alternative embodiment illustrated in  FIG.  13   , the portion  44  extends symmetrically to an axis C inclined with respect to the axis A and is arranged above the axis A with reference to the normal operating configuration of the helicopter  1 . 
     According to an alternative embodiment illustrated in  FIG.  14   , the portion  44  extends symmetrically to an axis D inclined with respect to the axis A and is arranged below the axis A with reference to the normal operating configuration of the helicopter  1 . 
     The support body  40  also comprises a heat dissipation device  100  provided to protect the heat exchanger  20  from possible damage caused by the heat transmitted by the motor system  6 . 
     In greater detail, the device  100  comprises ( FIG.  7   ):
         a grid  101  applied on the wall  65  in a position radially interposed between the heat exchanger  20  and the nozzle  15 ; and   a grid  102  interposed between the wall  66  and the wall  52 .       

     In particular, the grid  101  is shaped like an arc symmetrical with respect to the axis A and having a lower angular extension of the wall  65 . 
     The grid  101  is arranged below the walls  50 . 
     The grid  102  extends obliquely to axis A. 
     The helicopter  1  also comprises a device  55  for protecting the heat exchanger  20  from possible “heat shocks” which can temporarily overheat the oil present in the heat exchanger  20 . Said excess of heat can occur due to the hot gases, which therefore tend to stagnate to a small extent in the nozzle  15  or along the discharge duct  17 , once the motor system  6  is inactive. Another situation in which unpredicted overheating of the heat exchanger  20  can occur is the presence of flames inside the channels  28 , for example following a failure of the motor system  6  and and/or fire in the motor bay  8 . 
     In greater detail, the device  55  is selectively movable, between:
         an open configuration in which it allows the fluidic connection between the heat exchanger  20  and the discharge duct  17 ; and   a closed configuration in which it interrupts the fluidic connection between the heat exchanger  20  and the discharge duct  17 .       

     More precisely, the device  55  is arranged in the open configuration during the normal operation of the motor system  6  and/or in the absence of flames inside the channels  28 . 
     Conversely, the device  55  is arranged in the closed configuration when the motor system  6  is inactive or in the presence of flames inside the channels  28 . 
     In an embodiment of the invention, the device  55  is reversibly movable from the open configuration to the closed configuration through passive systems (for example elastic elements, shape memory metal alloys and the like) or through active systems (for example an electric, hydraulic or pneumatic actuator, or a suitable combination of the principles mentioned herein). 
     The helicopter  1  further comprises:
         a sensor  54  (only schematically illustrated in  FIG.  8   ) adapted to detect the fact that the motor system  6  is inactive and/or the presence of flames inside the channels  28 ; and   an actuator (not illustrated) controlled by the sensor and adapted to displace, according to what is detected by the sensor, the device  55  between the open and closed configurations.       

     In the case illustrated in  FIG.  8   , the device  55  is formed by the walls  50 , which are hinged to the support body  40  around a common axis parallel to the axis A and coinciding with the common ends  57  of the respective walls  50 . 
     The ends  59  leave the respective openings  43  free when the device  55  is in the closed configuration and leave said openings  43  free when the device  55  is in the open configuration. 
     In use, the first flow rate of air is sucked from the air intake  9  and reaches, through the intake duct, the compressor  11  of the motor system  6 . 
     The air intake  10  allows the entry of a stream of air into the motor bay  8 . 
     The first flow rate of air is compressed inside the compressor  11  and reacts with the second fuel flow rate inside the combustion chamber  13  generating the third flow rate of exhaust gases and air at high temperature and pressure. 
     Subsequently, the third flow rate of exhaust gases and air expands into the turbine  14  by driving the compressor  11  and the outlet shaft in rotation around the axis A. 
     Said third flow rate expands further into the nozzle  15  by reducing its static pressure at the downstream section  39 . 
     Said reduced static pressure at the downstream section  39  causes a fourth flow rate of air to be drawn through the air intake  25  and the ducts  26 . Said fourth flow rate reaches the openings  43  of the channels  28  in fluidic connection with the downstream section  39  of the nozzle  15 . 
     Said fourth flow rate of air, crossing the heat exchanger  20 , cools it and mixes with the third flow rate in the downstream section  39  of the nozzle  15 , so as to form the fifth flow rate. 
     The fifth flow rate of exhaust gases and air further expands in the nozzle  16 , therefore reducing its own static pressure at the downstream section  49  of the nozzle  16  itself. 
     Thanks to said reduction of the static pressure, the ejector  90  generates, at the downstream section  49 , a sixth flow rate of low temperature air inside the motor bay  8  and through the air intake  10 . 
     Said sixth flow rate of air by-passes the compressor  11 , the combustion chamber  13  and the turbine  14 , and cools the motor bay  8 . 
     Said sixth flow rate of air flows from the motor bay  8  to the downstream section  49  through the groove  81  and the openings  68  of the portion  45 , so as to cool the motor bay  8  ( FIG.  2   ). 
     Said sixth flow rate of air mixes with the fifth flow rate of air at the downstream section  49 , so as to form the seventh flow rate of air. 
     Said seventh flow rate of air crosses the discharge duct  17  until it reaches the mouth  18 , through which it is emitted into the atmosphere. 
     The device  55  is arranged in the open configuration during normal operation of the motor system  6  and/or in the absence of flames inside the channels  28 . 
     In said open configuration, the device  55  does not interfere with the flow of the stream of air in the channels  28 . 
     If the sensor  54  identifies that the motor system  6  is inactive or the presence of flames inside the channels  28 , the device  55  is arranged in the closed configuration, for example by means of a relative actuator, for example by rotation of the walls  50  around the common axis for hinging to the support body  40 . 
     In said closed configuration, the device  55  prevents the return of flames through the channels  28  and towards the heat exchanger  20 , preserving their integrity. 
     The device  100  favours the dissipation of the heat generated by the motor system  6  preferably within the motor bay  8 , further contributing to preserving the integrity of the heat exchanger  20 . 
     More precisely, the air heated by the motor system  6  rises inside the nozzle  15  until it reaches the grids  101 ,  102 , which allow it to escape and be disposed of. 
     From an examination of the characteristics of the helicopter  1  made according to the present invention, the advantages that it allows to obtain are evident. 
     In particular, the downstream section  39  of the nozzle  15  is fluidically connected with the heat exchanger  20  and the downstream section  49  of the nozzle  16  is fluidically connected with the motor bay  8 . 
     Consequently, the ejector  80  generates the fourth flow rate of air, which crosses and cools the heat exchanger  20 . 
     The ejector  90  generates the sixth flow rate of air, which crosses and cools the motor bay  8 . 
     Since the fourth and sixth flow rate of air travel through respective distinct paths upstream of the discharge duct  17 , it is possible to independently control the cooling of the heat exchanger  20  and the cooling of the motor bay  8 , unlike the solutions of the known type and described in the introductory part of the present description. 
     Consequently, a more precise and accurate control of the temperatures of the motor system  6  and/or of the motor bay  8  is possible without the installation of additional devices, which complicate the maintenance of the helicopter  1  and weigh on the overall weight of the same, as can be seen in the solutions of the known type and described in the introductory part of the present description. 
     The lobes  67  define an extension of the nozzles  15 ,  16 . In particular, the lobes  67  represent a completion from the fluid dynamic point of view of the interaction between the nozzles  15  and  16 . 
     In fact, as previously described, the lobes  67  are preferably arranged at the respective lobes  32  of the nozzle  15 , running parallel to the axis A. 
     Thanks to this, it is possible to reduce the turbulence and the fluid dynamic losses due to the flow of the third and fifth flow rate of air and exhaust gases inside the respective nozzles  15 ,  16 . 
     The device  100  allows to dissipate the heat generated by the operation of the motor system  6  inside the motor bay  8 , reducing the risk of damage to the heat exchanger  20 . 
     Heat dissipation occurs by convection and is made more efficient by the fact that the grid  101  is located above the motor system  6 . Thanks to said position, “hot” air and, hence, with a lower density of the air present in the motor bay  8 , present in the motor system  6 , naturally tends to move towards the grids  101 ,  102  and move away from the heat exchanger  20 . 
     The device  55  is selectively movable, below between:
         an open configuration in which it allows the fluidic connection between the heat exchanger  20  and the discharge duct  17  through the channels  28 ; and   a closed configuration in which it interrupts the fluidic connection between the heat exchanger  20  and the discharge duct  17  through the channels  28 .       

     In this way, it is possible to further reduce the risk of damaging the heat exchanger  20 , following the stagnation of hot air inside the channels  28  once the motor system  6  is inactive in the presence of open flames. 
     With detail in the accompanying  FIG.  12   , the portion  44  extends symmetrically to the axis B orthogonal to the axis A and is arranged above the axis A. Therefore, the inflow of fresh air through the air intake  25  benefits from a dynamic stream induced by the flow downwards and towards the air intake  25  generated by the main rotor  3 . Said dynamic stream is of the same order of magnitude as the fourth and sixth flow rates generated respectively by the ejectors  80 ,  90  through the heat exchanger  20  and the motor bay  8  respectively. 
     With particular reference to  FIG.  13   , the portion  44  extends symmetrically to the axis C inclined with respect to the axis A. In this condition, the contribution of the aforesaid dynamic stream is less than in the configuration of  FIG.  12    and the ejectors  80 ,  90  generate the predominant part of the fourth and sixth flow rate respectively through the heat exchanger  20  and the motor bay  8 , respectively. 
     With reference to  FIG.  14   , the portion  44  extends symmetrically to the axis D inclined with respect to the axis A and arranged below the axis A, with reference to the normal operating configuration of the helicopter  1 . 
     In said configuration the dynamic contribution of the rotor  3  is practically negligible. Therefore, the draw of the fourth flow rate of air and sixth flow rate of air is effectively obtained only through the respective ejectors  80 ,  90 . 
     Finally, it is clear that modifications and variations may be made to the helicopter  1  described above without thereby departing from the scope of protection of the present invention. 
     In particular, the helicopter  1  could comprise a pair of motor systems  6  having respective outlet shafts operatively connected to the main rotor  3 . 
     The aircraft capable of hovering could be a convertiplane instead of the helicopter  1 .