Abstract:
An aircraft ( 10 ) with improved aerodynamic performances is adapted to keep the directional stability and a very good aerodynamic behavior at medium-high incidence. The aircraft ( 10 ) includes a fuselage ( 12 ) to which shaped wings ( 18, 20 ) are associated, and a nose ( 52 ). The aircraft ( 10 ) also includes a vortex control device ( 72 ) of the extension of the leading edge of the wing at the root (LERX), shaped in order to symmetrize the bursting of the vortices generated by such LERX with a medium-high incidence. Said aircraft comprises removable equipment with at least one dissipation device of incident radar waves, on at least one hot portion of the aircraft.

Description:
This application is a National Stage Application of PCT/IB2011/001230, filed 30 May 2011, which claims benefit of Serial No. TO2011A000122, filed 14 Feb. 2011 in Italy and which applications are incorporated herein by reference. To the extent appropriate, a claim of priority is made to each of the above disclosed applications. 
     BACKGROUND 
     The present invention relates to a configuration of an aircraft having high aerodynamic performances and high performances with secondary operational capabilities. 
     Many kinds of aircraft must be easily piloted and must have particular dynamic characteristics in view of the missions to accomplish. 
     Such aircraft are often required to operate in airspace by flying at high incidence; this is in particular referred to the great incidence angle which the aircraft forms with respect to its own velocity vector, at each instant of time. 
     It is immediate to realize that, in such flight conditions, the aircraft must become extremely stable and easily controllable by the pilot, in order to keep a safe flight trim during the fighting phases. 
     Such stability is obtained by using particular automatic control apparatus, which permit the generation of forces and moments adapted to counterbalance undesired flight effects. 
     While the stability of the aircraft along the pitch axis can be suitably balanced through an optimized ratio between the position of the centre of gravity and the dimensions of the horizontal tail planes, the presence of lateral-directional instabilities (along the roll and yaw axis) with great incidence can be controllable with difficulty even adopting sophisticated automatic control apparatuses. 
     In this respect, it is therefore necessary to maximize the lateral-directional stability of the aircraft up to the great incidences, in order to increase the control possibility and the easy maneuvering so to prevent rapid and undesired deviation of the aircraft from a planned path. 
     Traditionally, and in particular, in recent days, attempts were made to obviate the instability drawbacks by acting on the aerodynamic profile of the fuselage and of other parts of the aircraft, but without reaching particularly valid results. 
     The use of aircraft on battlefields requires that they are not visible to the radar systems. 
     Aircraft of the known type are known with the term “stealth”, which are provided with a very sophisticated structure suitably studied for the present purpose, that is to be invisible to the radar systems. 
     The structure of the stealth aircraft greatly reduces the reflection towards the point of observation of the radiated electromagnetic waves, so making the aircraft substantially invisible to the radar systems. 
     Furthermore, such aircraft are completely painted with absorbing paints, which absorb the incident electromagnetic waves, so making the aircraft substantially invisible to the radar systems. 
     Such solution is costly for the realization of the aircraft itself, and its aerodynamic configuration is less than favorable so making the behaviour of the aircraft insufficient during the flight at high incidence. 
     SUMMARY 
     Within the aforementioned needs, one purpose of the present invention is therefore to obviate the cited drawbacks and in particular, to propose a configuration of aircraft with improved aerodynamic performances, which permits to optimize the behaviour of the aircraft mainly in the case of a flight with high incidence. 
     Another purpose of the present invention is to indicate a configuration of aircraft with improved aerodynamic performances, which permits to reduce the “buffet” effects characteristic of the wings with low elongation with a thin profile and variable centerline. 
     Further purpose of the present invention is to realize a configuration of aircraft with high aerodynamic performances, which permits to successfully prevent the loss of lateral-directional stability and the negative effects produced by the engine jet adjacent to the wall of the fuselage and the horizontal tailplane, in terms of resistance, stability and longitudinal control. 
     Further purpose of the present invention is to realize a configuration of aircraft with improved aerodynamic performances, making possible the exit from the spin, in general by optimizing the behaviour of the aircraft with high incidence. 
     Further purpose of the present invention is to realize a configuration of aircraft with high aerodynamic performances, which permits the installation of at least one battle equipment, as for example for the reduction of the radar signature, removable and adapted to make substantially the aircraft invisible to the radar systems. 
     In an advantageous way, the aircraft which is the subject of the present invention is designed, in particular, as an aircraft with high performances with secondary operational capabilities. 
     The configuration form is twin-engine and it is characterized by the presence of a series of extremely particular design features. 
     The preferably two-sided (tandem) cabin with mutually connected flight commands is first of all placed side by side to a nose having a substantially circular and variable cross-section, with a small elongation, which is optimized for a flight with high incidence, in which a radar can be integrated for the operating version. 
     The shape and dimensional characteristics of the nose are optimized in order to reduce the vortex interference of the same upon the aerodynamic characteristics of the aircraft with medium-high incidence; the mentioned characteristics furthermore permit to reduce the directional asymmetries with high incidence, which are typical of the traditional noses having a circular or elliptical cross-section. 
     Furthermore the wing profile is so modified, with respect to the profiles currently provided, that it integrates a system of minimization of the “buffet” effects characteristic of the wings with a small elongation and a thin profile, having a variable centerline. 
     The aerodynamic project further provides for positioning a control device of the vortices of LEX (LEX=“Leading Edge Extension”), suitably shaped in order to symmetrize the explosion of vortices generated by the LEX with a medium-high incidence, due to the fact that the symmetrical explosion of such vortices permits to keep the lateral-directional stability and the control of the aircraft with a medium-high incidence. 
     The training aircraft according to the invention has further an engine air intake adapted to guarantee the performances and the suitable fluid-dynamic interface with the engine; such project does not require the integration of a typical boundary layer on the upper side of an air intake integrated with a LEX. 
     The uncoupling of the horizontal and vertical tail plane permits to obtain a reduction of the aerodynamic resistance generated by the rear fuselage, to optimize the spin behaviour of the aircraft and improve the aerodynamic project of the same for the high incidence. 
     Finally, the removable equipment for the reduction of the radar signature of the aircraft is applicable in at least one warm portion of the aircraft, by keeping the aerodynamic characteristics of the aircraft (V). 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Further purposes and advantages of the present invention will be evident from the following description and the annexed drawings, given in a purely exemplary and non limitative way, in which: 
         FIG. 1  is a side view of an aircraft, in particular a training aircraft, designed according to the present invention; 
         FIG. 2  is a top view of an aircraft, in particular a training aircraft, designed according to the present invention; 
         FIG. 3  is a bottom view of an aircraft, in particular a training aircraft, designed according to the present invention; 
         FIG. 4  is a front view of an aircraft, in particular a training aircraft, designed according to the invention; 
         FIG. 5  is a rear view of an aircraft, in particular a training aircraft, designed according to the invention; 
         FIG. 6  is a cross-sectional view, taken along line VI-VI of  FIG. 2 ; 
         FIG. 7  is a partial and enlarged view of a detail of the configuration of aircraft designed according to the present invention; 
         FIG. 8  is a cross-sectional view, taken along the line VIII-VIII of  FIG. 7 ; 
         FIG. 9  is a cross-sectional view, taken along the line IX-IX of  FIG. 7 ; 
         FIG. 10  is a cross-sectional view, taken along the line X-X of  FIG. 7 ; 
         FIG. 11  is a cross-sectional view, taken along the line XI-XI of  FIG. 7 ; 
         FIG. 12  is a cross-sectional view, taken along the line XII-XII of  FIG. 7 ; 
         FIG. 13  is a cross-sectional view, taken along the line XIII-XIII of  FIG. 7 ; 
         FIG. 14  is a cross-sectional view, taken along the line XIV-XIV of  FIG. 7 ; 
         FIG. 15  is a cross-sectional view, taken along the line XV-XV of  FIG. 7 ; 
         FIG. 16  is a cross-sectional view, taken along the line XVI-XVI of  FIG. 7 ; 
         FIG. 17  is a cross-sectional view, taken along the line XVII-XVII of  FIG. 7 ; 
         FIG. 18  is a perspective and enlarged view of a detail of the configuration of aircraft designed according to the present invention; 
         FIGS. 19A ,  19 B and  19 C show an aircraft, according to the present invention, in different projected views in which the hot portions are evidenced, which typically are the most significant and can be made invisible by the radar system through the reduction equipment of the radar signature; 
         FIG. 20  shows in cross-section a metallization made to the transparent surfaces of the aircraft, according to the present invention; 
         FIGS. 21A and 21B  show the cover portion adapted to make the attaching edges of an aircraft invisible to the radar systems, according to the present invention,  FIG. 21A  respectively showing a cross-section in a perspective vision,  FIG. 21B  showing a lateral cross-section of the cover portion; 
         FIGS. 22A and 22B  show the equipment for the first ordinate of the fuselage,  FIG. 22A  respectively showing the position of the equipment;  FIG. 22B  showing a portion of the cover used for making the first ordinate of the fuselage of the aircraft invisible to the radar systems, according to the present invention in a front cross-sectional vision; 
         FIGS. 23A ,  23 B and  23 C show the grid adapted to make the engine face of the aircraft invisible to the radar systems, according to the present invention; in particular,  FIG. 23A  shows the grid in a perspective vision,  FIG. 23B  shows a detail of the cross-section of a grid,  FIG. 23C  shows the substitution structure to be substituted to such grid once having been removed. 
     
    
    
     DETAILED DESCRIPTION 
     With reference to the cited figures, an aircraft is shown, in particular a training aircraft, having a configuration with improved aerodynamic performances, according to the present invention, generally indicated with the reference number  10 . 
     Aircraft  10  comprises a fuselage  12 , having an upper wall  14  and a lower wall  16 , and two wings, one right wing  18  and one left wing  20  respectively, being connected with fuselage  12 . 
     Right wing  18  has a wing end  22 , whereas left wing  20  has end  24 . 
     Aircraft  10  further comprises a directing rudder  34 , mounted on tail drift  8  or vertical tail plane and a horizontal tail plane  44 , having one right horizontal stabilizer  26  and one left horizontal stabilizer  28 , with respective ends  30 ,  32 . 
     In preferred but not limitative embodiments of the invention, as previously noted, the typical configuration shape is one twin-engine having two air intakes  46  for the inlet of corresponding turbo-jets  48 , having respective outlet cones  60 . 
     Finally, near nose  52 , in which a radar can be integrated for the operative version of aircraft  10 , a cockpit  54  is installed, preferably two-sided (tandem) with mutually connected flight commands, which is protected by a windshield  62  and furthermore a probe  58  can also be provided, for the operation of refueling aircraft  10  in flight. 
     With particular reference to  FIGS. 2 and 3 , each wing  18 ,  20  of aircraft  10  has outer ailerons  56  and inner flaps for take-off and landing  64 , with a double slot, which are provided at the rear profile or trailing edge  70  of each wing  18 ,  20 , and further optimizing devices of the wing maneuvering profile or mobile leading edge (leading edge droops)  66 , present at leading edge  68  and the profile of which is shaped according to a particular geometry, on the basis of the general aerodynamic considerations cited in the present description. 
     More particularly, the technical features of aircraft  10 , adapted to obtain high aerodynamic performances and flight stability, according to the present invention, are the following. 
     First of all, the aerodynamic project is characterized by the presence of a control device (LVC=“LEX Vortex Controller”) of the vortices of the LEX (“Leading Edge Extension) with a medium-high incidence (reference number  72  of  FIG. 1 ). 
     In fact, the presence of LEX, having the plan gothic shape equal to 6.4% of the gross wing surface (as in the case of the present invention), permits to generate a vortex lift with high incidence and the project of the LEX is further refined with the integration of a LVC (“LEX Vortex Controller”) at the end of the same, in order to assure the symmetrical explosion of the vortices with high incidence with yaw attitudes, so preventing the consequent loss of lateral-directional stability. 
     The dimensions of control device  72  depend on those of the LEX facing it, and in any case, the greater the LEX, the higher the LVC; the tolerance can be defined in terms of the ratio between the surface of one LEX and the height of the corresponding LVC, according to which the design value of this ratio is 2.35 m and the tolerance field to apply varies between +100% and −50% with respect to the design value. 
     The shape of nose  52  of aircraft  10  and its dimensional features are further optimized in order to reduce the vortex interference of the same on the aerodynamic features of aircraft  10  with a medium-high incidence; the cited features further permit to reduce the directional asymmetries with high incidence, which are typical of the noses having a traditional circular or elliptical cross-section. 
     Nose  52  (“forebody”) of aircraft  10  according to the invention has a series of cross-sections having a different geometry starting from tip  74  up to a connecting corner with the apex of the LEX. 
     An exemplary and preferred, but not limitative, embodiment of the geometric shape and of the successive cross-sections having a different geometry, between tip  74  and a reference cross-section taken at an altitude  76  (substantially positioned at the beginning of the cockpit  54 ), is illustrated in sequence in  FIGS. 8-17 , from which it can be derived that from a substantially circular cross-section with a low extension ( FIGS. 8-11 ) one goes to a cross-section having an oval geometrical or dome shape ( FIGS. 12-17 ). 
     From the mentioned figures also the offset position of nose  52  can be seen, from tip  74  up to the reference cross-section illustrated in  FIG. 17 , with respect to longitudinal axis K. 
     In particular, according to a preferred embodiment of the invention, the ratio between the length of nose  52 , taken from tip  74  up to the cross-section made along the line XVII-XVII (reference L), and the average between the lengths A and B of the two half-axes of the cross-section of the same (cross-section illustrated in  FIG. 17 ) has a value of 1.873, with a tolerance of ±10%. 
     The special construction and its reflection in terms of the aerodynamic conditions in flight just derives from the combination of the parameter mentioned before (more or less the eventual tolerance) with the evolution of the cross-sections of nose  52 , from the apex or tip  74  of aircraft  10  up to the reference cross-section taken along the line XVII-XVII. 
     In  FIG. 18  an engine air intake is also shown in detail, generally indicated with  46 , which contributes to guarantee the performances of aircraft  10 , mainly regarding the adequate fluid-dynamic interface with the relative turbojet engine. 
     Air intake  46  has a radius at the evolutive leading edge, optimized in order to reduce the distortion to engine face  47  on the inner side, due to the high incidence on the lower portion and for reducing the bleed resistance on the outer side. 
     In particular, the average radius at the leading edge of inner lip  76 A has a value of 7 mm, whereas the average radius of lower lip  78  is 17.5 mm and that of outer lip  80  is 14 mm, so that the capture area of the air intake is approximately 0.322 m 2 , the throat area of the air intake is approximately 0.257 m 2  and the inlet area at the engine is approximately 0.273 m 2  (it must be remembered that such dimensions are referred to an air intake). 
     Air intake  46  is characterized by the absence of a typical spacer of the boundary layer (“diverter”) on the upper side of each intake integrated with one LEX, thanks to the particular relationship between the length of the LEX and its shape itself; in fact, the LEX acts as a flow straightening shield with a high incidence. 
     The air intake system can further provide for the presence of two additional intakes (non illustrated in the figures), positioned on the back of the connection between wing  18 ,  20  and fuselage  12 , which open when the pressure in the connection is lower than the pressure on the back of the connection wing-fuselage, thanks to preloaded springs integrated in the hinge of the same additional intakes. 
     The function of such air intakes is to reduce, when opening, the local incidences on the lips of the main air intake  46  with a high incidence, so reducing the quantity of air passing through the cited main air intake  46 . 
     One of the particular features of aircraft  10 , which guarantee its high performances with respect to its flight stability and aerodynamic structure, is surely made by the uncoupling of horizontal  44  and vertical tail plane  38 , which permits to reduce the aerodynamic resistance generated by the rear fuselage, optimizing the spin behaviour of aircraft  10  and improve the entire aerodynamic project of the same for its high incidence. 
     The vertical empennage with a trapezoidal shape comprises a rudder  34  and is coupled with the wing, in the sense that the leading edge of the same, generally indicated with  36  in  FIG. 1 , permeates trailing edges  70  of each wing  18 ,  20 , in order to make it possible the exit from the spin and for optimizing in general the behavior of vehicle  10  with a high incidence. 
     The horizontal empennage, also characterized by a trapezoidal shape, is actuated by two independent actuators, which permit the symmetrical and asymmetrical deflection of the same; such empennage has finally a hinge axis, indicated with  86  in  FIG. 2 , which is inclined rightwards and leftwards about 7.5°, with respect to a transversal axis  88 , in order to optimize the inertia and hinge moments. 
     The uncoupling between the horizontal and vertical tail plane can further be characterized by defining a tolerance on a reference parameter, defined as the ratio between altitude C, shown in  FIG. 1  and which can be defined as the distance between the apex at the root of the drift and the apex at the root of horizontal tail plane  44 , and the tail arm, being of 4181 mm; it follows that the reference value cited before is 1932 mm/4181 mm=0.462, with an applicable tolerance of 10%. 
     Even the wing profile is modified and optimized, with respect to the training aircraft of the traditional kind, in order to reduce the “buffet” effect, by considering the known features of a wing having a low elongation with a thin profile and variable centerline near the saw tooth. 
     According to the invention, on the contrary, a wing is used (references  18  and  20 ) having a trapezoidal shape, with a medium elongation (AR=4), characterized by the presence of a saw tooth (indicated with S in  FIG. 2 ) being 67.5% of the gross wing aperture; the modification with respect to the traditional wings relates first of all to the radius of the leading edge, indicated with R in  FIG. 6 , the shape of which varies from the circular one (known art) to the triangular one, in order to optimize the position of the stagnation point, in the presence of leading edge  68  and of “Leading Edge Droops”  66  deflected at medium incidences. 
     As can be clearly seen from  FIG. 6 , which shows an enlarged cross-section along line VI-VI of  FIG. 2 , each wing  18 ,  20  is provided with a profile with a variable curvature, both at leading edge  66  (“Leading Edge Droop”) and at trailing edge  70 , by ailerons  56 ; these are programmed only in the transonic field, in order to provide for a reduction of the curvature which alleviates the compressibility effects. 
     In quantitative terms, the design value of the cord percentage extension at the leading edge is 0.36%, with a tolerance from +0.5% and −0.2%, with respect to the nominal value, whereas the design value of the gross percentage wing aperture at the modification of the profile, with respect to the traditional solutions, is 8.2%, with a tolerance from +10% and −5%, with respect to the nominal value. 
     Further features of aircraft  10  are represented in fuselage  12 , which in its rear portion  16 , provides for the integration of the engine outlets and the presence of a small stern, indicated with  90  in  FIG. 3 , which supports the tail planes. 
     Also the region corresponding to the engine outlets is optimized in order to reduce the negative effects, in terms of resistance and stability/longitudinal control, produced by the engine jet adjacent to fuselage wall  12  and horizontal empennage  44 . 
     Vehicle  10  is further made of a tricycle cart, comprising one front and two main carts, the front cart of which has a stem, with four closing doors of the space, and with a feedback towards the flow direction. 
     The main cart retracts in an opposite direction with respect to the flow direction and the feedback system is optimized in order to permit to install outer ventral loads into the fuselage. 
     Aircraft  10  according to the present invention integrates an automatic flight control system (“Fly By Wire”), of the digital quadruplex kind, which permits to optimize the performances and the flight qualities; the system then permits to improve the flight safety through the automatic limitation of the flight regimes, which could be non comfortable for the pilot or could cause the loss of control (“Carefree Handling”). 
     The aircraft according to the present invention is provided with an equipment for the reduction of the radar signature of an aircraft, in at least one hot portion “H” of the aircraft itself, which is easily detectable by the radar systems. Such equipment comprises at least one device for dissipating incident radar waves, which can be applied and subsequently removed, in function of the needs, always maintaining the aerodynamic features of the aircraft. 
     For the purposes of the present invention, a hot “H” portion of the aircraft is defined as any of the portions normally detectable by a radar system, like for example: a cockpit  54 , comprising transparent portion (canopy and windshield)  62 ; one first ordinate of fuselage  12 , with which nose  52  is connected, to which the radar antenna of the aircraft is bound, visible from the radar through the nose made of radome transparent to the radiation itself; a plurality of leading edges ( 36 ,  66 ,  68 ) of the components like for example wings ( 18 ,  20 ), engine air intakes  46 , fog tails ( 38 ,  44 ) and at least one engine face  47 . 
     Such equipment provides for at least one device, with a particular or respective technical solution, for each hot portion of the aircraft, in order to reduce the radar signature. 
     For reducing the radar signature coming from the hot portions “H” of the aircraft like a cockpit  54 , comprising transparent portion  62 , the equipment comprises at least one metallization  100 , which is made on transparent portion  62 . 
     Such metallization  100  is adapted to restore an electrical continuity of the aircraft, so reducing the generation of diffractions of the incident wave, generated within the cockpit  54  covered by windshield and potentially receivable by the radar system; furthermore, it assures reflections outside the coverage of the radar system. 
     Metallization  100  is realized through the application of a plurality of coating layers, preferably three layers. 
     In the embodiment shown in  FIG. 2  such metallization  100  comprises at least one first layer or base  101 , which is able to prepare the windshield which must receive metallization  100 . 
     The deposit of such first layer  101  is followed by at least a second layer  102 , preferably through the deposit of material having a high degree of electrical conductivity, like for example gold or equivalent materials with high capacity of being mold on the surface. Such second layer  102 , which in fact represents the conductive metalizing layer, is applied upon first layer  101  through atomization methods of the material. 
     For the protection of metallization  100  at least one protective coating  103  is deposed, adapted to minimize the risks of damages of metallization  100 , due to accidental shocks or atmospheric agents. 
     In addition to the plurality of layers cited above, metallization  100  comprises a plurality of electrical devices adapted to guarantee the electrical connection of metallization  100  to the structure of the aircraft. 
     In the embodiment illustrated in  FIG. 20  such electrical device comprises at least one junction device  105 , preferably realized through a conductive sheet, for example of silver, being in electrical contact with second layer  102 . 
     Such junction device  105  is able to connect the layers comprised in metallization  100  with the structure of the aircraft. 
     Such metallization  100  can be realized together with the realization of the transparent portions of the cockpit, which can be completely substituted. 
     Once having finished the use of the equipment on the aircraft it is sufficient to substitute windscreen  62  of the cockpit with transparent portions in which metallization  100  is absent. 
     For reducing the radar signature coming from the hot “H” portions of the aircraft, such for example at least one leading edge ( 36 ,  66 ,  68 ,  70 ) of the components like wings ( 18 ,  20 ), the equipment comprises at least one cover portion  200 , positioned on the front edge of such components of the aircraft, by maintaining the aerodynamic profile of the component itself. 
     In the embodiment illustrated in  FIGS. 21A and 21B  such cover portion  200  comprises a first support structure  202  made of a preferably metal material, fixed to the structure of the component of the aircraft through fastening means, like screws or bolts. 
     Upon such first structure  202  a second radar-absorbing structure  203  is bound, adapted to absorb the incident electromagnetic waves so greatly attenuating the eventually reflected and/or refracted waves. 
     Once having terminated the use of such cover portion  200  it is unhooked from the structure of the aircraft itself and possibly substituted with a cover portion which however keeps its aerodynamic profile through a suitable shaping of the structure  202 , upon which the second radar-absorbing coverage  203  is absent. 
     Such solution permits to keep the aerodynamic profile required for such leading edges, ( 36 ,  66 ,  68 ,  70 ), so reducing the costs and the complexity of installation. 
     For reducing the radar signature coming from the hot “H” portion of the aircraft, like the first ordinate of fuselage  12 , at nose  52  the equipment comprises at least one sheet of adhesive metal material  301 , for example of aluminum, which covers the first ordinate of fuselage  12  of the aircraft and at least one layer of absorbent material  302 , fixed to such sheets of metal material  301 . 
     As shown in  FIG. 22A  the equipment is positioned between the first ordinate of fuselage  12  and radar antenna “A” which is coated by radome nose  52 . 
     From a front vision, the shape of the sheets of metal material  301 , and of absorbing material  302 , are such to correspond to the shape of the first ordinate of fuselage  12  of the aircraft, with the features cited before. 
     As shown in  FIG. 22B  the absorbent material  302  is fixed for example by means of glue, to the sheets of metal material  301 . 
     Absorbing material  302  used is for example a sponge layer soaked with ferrite powder, graphite etc., adapted to absorb the incident electromagnetic waves so greatly attenuating the reflected wave. 
     Once having terminated the function of the equipment for reducing the signature of the first ordinate of fuselage  12  of the aircraft, this is directly done by removing the sheets of metal material  301 , to which layers  302  are fixed, so restoring the first ordinate of fuselage  12  back to the original state. 
     Such solution permits to reduce the costs and the weight of the equipment to be fastened to the aircraft, in addition to reducing the complexity for the fastening. 
     For reducing the radar signature coming from the hot portions of the aircraft, like at least one engine face  47 , the equipment comprises at least one grid  400  adapted to permit the entry of an air flow towards the engine and reduce the visibility of engine face  47  of the aircraft by the radar system. 
     In the embodiment shown in  FIGS. 23A and 23B  grid  400  comprises an internal structure in which a plurality of apertures  401  are realized, having such dimensions to be like a plane surface for the lowest frequencies of the frequency spectrum, commonly used in the radar recognition systems operating at a low frequency, like for example Jet Engine Modulation, so avoiding the generation of diffracted waves which can be recognized by the radar system itself. 
     The internal surface of such apertures  401  is coated with a radar-absorbing material of a small thickness adapted to absorb the electromagnetic waves at higher frequencies, like for example in a X-band around 10 GHz. 
     The combination of the dimensions of apertures  401  and of the absorbing material, permits to make such component of the aircraft detectable with difficulty by the radar systems. 
     Preferably, such grid  400  has a circular shape similar to the cross-section of the structure of the engine compartment. 
     In the embodiment shown in  FIGS. 23A ,  23 B grid  400  comprises a support structure  402 , comprising a plurality of rings, adapted to fix such grid  400  to the aircraft through fastening means, like for example screws or bolts. 
     In the detail of  FIG. 23B , support structure  402  comprises a first ring  403 , which will be structurally bound to the duct of air intake  46  of the engine, a second ring  404 , which will be bound to a fireproof bulkhead comprised in the engine compartment and a third ring  405 , adapted to block the gasket interfaced with the engine. 
     Once having terminated the time in which the use of such grid  400  is necessary, it can be extracted from the structure of the engine compartment, and possibly it can be substituted, in order to keep the continuity of the duct, with a substituting structure  406 , which essentially has the outer shape of the support structure and in which the inner structure of grid  400  itself is absent. 
     The equipment with which the aircraft is provided, according to the present invention, permits to obtain cost advantages both in the phase of realization and in the phase of maintenance, as it is only used when such application is necessary, so reducing the wear of such equipment. 
     The equipment only acts on the most important hot regions so avoiding to waste resources in order to make portions of the aircraft detectable with difficulty by the radar systems, which are still detectable with difficulty themselves, so greatly reducing the costs of the equipment itself. 
     The choice of only intervening on the centers which can be more greatly detectable is an optimization point between the cost of the operations and the effectiveness of the obtainable benefits. 
     Such equipment preferably is applied to all hot portions of the aircraft, cited according to the present invention, but in some cases it is possible to apply the equipment just on some of such “H” portions, by keeping other ones in their initial configuration. 
     From the description made the features are clear of the configuration of an aircraft with improved aerodynamic performances, which is the subject of the present invention, and also clear are its advantages. 
     It is finally clear that numerous other variations can be made to the structure of aircraft in question, without for this reason abandon the novelty principles inherent in the inventive idea, and it is also clear that, in the practical implementation of the invention, the materials, the shapes and the dimensions of the details shown can be of any kind according to the needs and the same can be substituted with other technically equivalent ones.