Abstract:
An aircraft ( 10 ), in particular a trainer aircraft with improved aerodynamic performance, having a configuration able to keep a directional stability and a very good aerodynamic behaviour even at very high angles of attack, where, traditional configurations prove themselves inefficient. In particular, this configuration foresees a forebody ( 52 ) with variable section, optimised for high angle of attack flights, a LEX vortex control device ( 72 ) at least one diverterless air intake ( 46 ), and a wing profile ( 18, 20 ) optimised in order to reduce the buffet effects typical of low aspect ratio wings with thin profile and variable camber. The aircraft ( 10 ) presents, finally, staggered tails ( 44  and  38 ), to optimise the aerodynamic performance.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
   Not Applicable 
   STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
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   INCORPORATION-BY-REFERENCE OF MATERIAL SUBMITTED ON A COMPACT DISK 
   Not Applicable 
   REFERENCE TO A MICROFICHE APPENDIX 
   Not Applicable 
   BACKGROUND OF THE INVENTION 
   (1) Field of the Invention 
   The present invention is related to an aircraft configuration with high aerodynamic performance; in particular the aircraft according to this invention is designed like a high performance trainer with secondary operational capabilities. 
   (2) Description of Related Art 
   Many aircraft have to be easily flown and must have peculiar dynamic characteristics in accordance with the mission to be carried out. Typical examples of the above include light aircraft for aerobatics, trainers, and combat aircraft. 
   These aircraft often have to fly at a high angle of attack (the angle between the aircraft and wind speed direction at any time). It is easy to understand that, under these flight conditions, the aircraft has to be very stable and easily controlled by the pilot, in order to maintain a safe attitude during flight. This stability is obtained using special automatic equipment able to generate forces and moments to counterbalance undesirable flight attitudes. 
   While the aircraft stability along the pitch axis can be controlled optimizing the static margin and time-to-double amplitude, the presence of lateral-directional instability (on “roll” and “yaw” axes) at high angles of attack can be difficult to control even using highly sophisticated Flight Control Systems. It is therefore necessary to maximize the aircraft lateral-directional stability up to high angles of attack in order to allow aircraft control/agility and to avoid flight departures and spins. 
   Usually, and in particular lately, attempts have been made by simply modifying the aerodynamic shape of the fuselage and other aircraft parts. Up to now those attempts have not led to successful results. 
   In the frame of the requirements listed above, one of the goals of the present invention is, therefore, to avoid the mentioned problems and, in particular, the problem relevant to an aircraft configuration with improved aerodynamic performance, able to optimize the aircraft behavior especially during high angle of attack flights. 
   Another goal of the present invention is to present an aircraft configuration with improved aerodynamic performance, able to reduce buffet effects typical of low aspect ratio wings with thin profile and variable camber. 
   Another goal of the present invention is to realize an aircraft configuration with high aerodynamic performance, able to avoid successfully the loss of lateral-directional stability and the negative effects generated by the engine flow next to the fuselage side wall and the horizontal tail, as far as drag, stability and longitudinal control are concerned. 
   An additional goal of the present invention is to realize an aircraft configuration with improved aerodynamic performance, able to recover from spin, optimizing, in general, the aircraft behavior at high angles of attack. 
   BRIEF SUMMARY OF THE INVENTION 
   These and other goals are met by an aircraft configuration with improved aerodynamic performance wherein the aircraft has a fuselage ( 12 ) fitted with wings ( 18 ,  20 ), at least one air intake ( 46 ) and a fuselage forebody ( 52 ) with a varying profile starting from the apex ( 74 ) up to the edge coupling which is a roughly circular section to the LEX apex and at the LEX apex a chinned section, said aircraft ( 10 ) also having a LEX vortex control device ( 72 ) at mid/high angles of attack and staggered tails ( 44  and  38 ), obtained by coupling the vertical fin to the wing ( 18 ,  20 ), whose leading edges ( 36 ) overlap the tailing edges ( 70 ) of each wing ( 18 ,  20 ), in order to optimize the aerodynamic performance. 
   Preferably, the subject aircraft of the present invention is designed, in particular, like a trainer with high performance and secondary operational capabilities. The configuration includes a twin engine “formula” and it is characterized by the presence of a range of very peculiar structural details. 
   The twin-seat cabin (in tandem) with interconnected flight controls is coupled with a forebody ( 52 ) with basically circular and variable section, characterized by low aspect ratio, optimized for high angles of attack flights. In this space, radar for the aircraft operational version can be easily integrated. 
   The forebody shape and dimensions are optimized in order to reduce its vortex interference on the aircraft aerodynamic characteristics at mid/high angles of attack. These characteristics reduce the directional asymmetries at high angles of attack, typical of forebody with circular or elliptical section. 
   Moreover the wing profile of an aircraft configured in accordance with the present invention is different from the standard wing profiles, in order to integrate a system able to minimize the buffet effects, typical of low aspect ratio wings with thin profile and variable camber. 
   The aerodynamic design also includes a LEX (Leading Edge Extension) vortex control device, properly sized, in order to make symmetrical the LEX vortex bursting at mid/high angles of attack. This symmetric vortex bursting allows the lateral-directional stability and the aircraft control to be maintained at mid/high angles of attack. 
   The training aircraft under this patent has at least one engine air intake able to guarantee the performance and a proper fluid-dynamic interface with the engine. The design does not foresee the integration of a typical diverter on the upper lip of an air intake, integrated with a LEX. 
   Finally, the horizontal tail staggering allows for a decrease in the aerodynamic drag produced by the fuselage afterbody, to optimize the aircraft spin behavior and to improve the aircraft aerodynamic design for high angles of attack maneuvers. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
     Further goals and benefits of this patent shall be clear from the following description and from the attached drawings, provided solely to illustrate the present invention without limiting the invention in any manner, as follows: 
       FIG. 1  is a side view of an aircraft, in particular of a trainer aircraft, designed according to the present invention; 
       FIG. 2  is a top view of an aircraft, in particular of a trainer aircraft, designed according to the present invention; 
       FIG. 3  is a bottom view of an aircraft, in particular of a trainer aircraft, designed according to the present invention; 
       FIG. 4  is a front view of an aircraft, in particular of a trainer aircraft, designed according to the present invention; 
       FIG. 5  is a back view of an aircraft, in particular of a trainer aircraft, designed according to the present invention; 
       FIG. 6  is a section along the VI-VI line of  FIG. 2 ; 
       FIG. 7  is a partial and enlarged view of a detail of the aircraft configuration designed according to the present invention; 
       FIG. 8  is a section along the VIII-VIII line of  FIG. 7 ; 
       FIG. 9  is a section along IX-IX line of  FIG. 7 ; 
       FIG. 10  is a section along X-X line of  FIG. 7 ; 
       FIG. 11  is a section along XI-XI line of  FIG. 7 ; 
       FIG. 12  is a section along XII-XII line of  FIG. 7 ; 
       FIG. 13  is a section along XIII-XIII line of  FIG. 7 ; 
       FIG. 14  is a section along XIV-XIV line of  FIG. 7 ; 
       FIG. 15  is a section along XV-XV line of  FIG. 7 ; 
       FIG. 16  is a section along XVI-XVI line of  FIG. 7 ; 
       FIG. 17  is a section along XVII-XVII line of  FIG. 7 ; 
       FIG. 18  is an enlarged isometric view of a detail of the aircraft configuration designed according to the present invention. 
   

   DETAILED DESCRIPTION 
   Making reference to the attached drawing Figures, an aircraft, in particular a trainer aircraft, having a configuration with improved aerodynamic performance, accordingly to the present invention, is generally indicated with reference  10 . The aircraft  10  includes a fuselage  12 , having an upper side wall  14  and a lower side wall  16 , and two wings, respectively a right wing  18  and a left wing  20 , fitted to the fuselage  12 . 
   The right wing  18  has the tip  22 , while the left wing  20  has the tip  24 . The aircraft  10  features also a rudder  34 , fitted to the vertical fin  38  and a horizontal stabilizer  44  having a right horizontal stabilizer  26  and a left horizontal stabilizer  28  with respective tips  30 ,  32 . 
   In the preferred embodiment, but not limiting the present invention, as already mentioned above, the configuration shown in the drawing Figures is a twin-engine design which includes two air intakes  46  for the engines  48 , with relative engine nozzles  60 . 
   Finally, radar for the operational version of the aircraft  10  could be integrated in the fuselage forebody area  52 , with a two in tandem pilot cockpit  54  having interconnected flight controls, protected by a windshield  62 . A probe  58  could also be included, in order to carry out in-flight refueling for the aircraft  10 . 
   With reference to  FIGS. 2 and 3 , each wing  18 ,  20  of the aircraft  10  includes outer ailerons  56  and take-off and landing double slotted flaps  64 , integrated in the trailing edge  70  of each wing  18 ,  20 . Other devices for wing camber optimization, including leading edge droops  66 , are integrated in the wing leading edge  68 . Their profile is shaped following a particular geometry, on the basis of the general aerodynamic design mentioned in this description. 
   In particular, the technical characteristics of the aircraft  10 , aimed to obtain high aerodynamic performance and flight stability, according to the present invention, are as follows. 
   First, the aerodynamic design is characterized by the presence of an LVC (“LEX Vortex Controller”) device to control the LEX (“Leading Edge Extension”) vortex at mid/high angles of attack ( FIG. 1 , reference character  72 ). The LEX, with a gothic platform equal to 6.4% of the reference wing gross area (in accordance with the present invention), allows vortex lift generation at high angles of attack. LEX design is further refined integrating the LVC at its tip, in order to ensure the vortex symmetric bursting at high angles of attack for sideslip attitudes and to prevent loss of lateral-directional stability. 
   The size of the LEX vortex controller  72  depends on the size of the front LEX. In any case, the bigger the LEX is, the higher the LVC must be; the tolerance can be defined by the ratio between the LEX area and the height of the corresponding LVC. The design point of this ratio is equal to 2.35 m and the applicable tolerance range varies from 100% greater than, or 200% of, to 50% less than, or 50% of, the ratio from the design point. 
   The shape of the fuselage forebody  52  of the aircraft  10  and its dimensions are even further optimized in order to reduce its vortex interference on the aerodynamic characteristics of the aircraft  10  at mid/high angles of attack. The characteristics also allow the directional asymmetries at high angles of attack to be reduced, typical for the forebody with standard circular or elliptical sections. 
   The forebody  52  of the aircraft  10 , according to the present invention, shows a range of forebody sections with different geometry starting from the apex  74  up to the forebody edge merging with the LEX apex. 
   Exemplary and preferred, but not limiting, geometric shapes of the sequence of sections between the apex  74  and a section referred to the quote  76  (positioned more or less at the beginning of the twin-seat cabin  54 ), is presented in sequence at  FIGS. 8-17 . From these Figures it can be inferred that from a roughly low aspect ratio circular section ( FIGS. 8-11 ) a chined section is provided ( FIGS. 12-17 ). 
   The mentioned drawing Figures make clear the shift of the forebody  52  from the longitudinal axis K, from the apex  74  up to the reference section presented in  FIG. 17 . In particular, according to a preferred embodiment of the present invention, the ratio between the length of the forebody  52 , starting from the apex  74  up to the section along the XVII-XVII line (reference L), and the average between the lengths A and B of the two axes of the section (section shown in  FIG. 17 ), presents a value of 1.873, with a tolerance equal to ±10%. 
   The structural peculiarity and its effect on the flight conditions come out from the combination of the parameter mentioned above (plus or minus the tolerance, if any) and the sequence of the forebody  52  sections, from the apex  74  of the aircraft  10  up to the reference section along XVII-XVII line. 
     FIG. 18  shows in detail an engine air intake, indicated as item  46 , which contributes to guarantee the performance of aircraft  10 , above all as far as the fluid-dynamic interface with the relevant jet engine is concerned. 
   The air intake  46  shows a variable leading edge radius optimized in the lower part in order to reduce the engine face flow distortion at high angles of attack and in the side part in order to reduce the transonic spillage drag. 
   In particular, the inner lip average leading edge radius  76 A is equal to 7 mm, while the lower lip average leading edge radius  78  is equal to 17.5 mm and that of the outer lip  80  is equal to 14 mm. Thus, in the preferred embodiment the capture area of the air intake  46  is about 0.322 m 2 , and the throat area of the air intake  46  is about 0.257 m 2 , with the engine face area being about 0.273 m 2  (these values referring to one air intake  46 ). 
   The air intake  46  is diverter-less on the upper side of each air intake side and is integrated with a LEX, thanks to the peculiar ratio between LEX length and shape. The LEX effectively acts as a shield at high angles of attack. 
   The air intake  46  can also include two additional blow indoors (not showed in the Figures), positioned on the upper wing-body junction between wings  18 ,  20  and fuselage  12 , which open when the pressure in the duct is lower than pressure on the upper wing-body junction. The blow indoors use pre-loaded springs integrated in the blow indoor hinges. The blow indoors hinges aim to reduce, when they open, the local angles of attack on the air intake lips  46  at high angles of attack, reducing the air mass flow quantity passing through the above mentioned air intake  46 . 
   One of the characteristics that guarantee the high performance, the stability and the aerodynamic structure of the aircraft  10  is the staggered tails  44  and  38 . They reduce the aerodynamic drag generated by the fuselage afterbody, to optimize the spin behavior of the aircraft  10  and to improve its whole aerodynamic design for high angles of attack flights. 
   The vertical fin  38  with trapezoidal platform includes the rudder  34  and is coupled to the wing. This means that the leading edge of the vertical fin  38 , indicated by reference character  36  in  FIG. 1 , overlaps the trailing edges  70  of each wing  18 ,  20 . This configuration allows for the recovery from spin and to optimize the behavior of the aircraft  10  at high angles of attack. 
   The horizontal stabilizer  44  with trapezoidal platform is moved by two independent actuators which allow its symmetric and asymmetric deflection. The horizontal stabilizer  44  presents a hinge axis, indicated by reference character  86  in  FIG. 2 , which is oriented, on the right and on the left, at about 7.5° from a lateral axis  88 , in order to optimize the hinge and the inertial moments. 
   The staggered tails can be further characterized by a tolerance referred to as the ratio between quote C, as shown in  FIG. 1  and defined as the distance between the apex of the vertical fin root chord and the apex of the horizontal tail root chord  44 , and the tail arm, equal to 4181 mm. It follows that the reference ratio given above is equal to 1932/4181 mm, or 0.462, with an applicable tolerance equal to 10%. 
   Also, the wing profile is modified and optimized in comparison with traditional trainer aircraft, in order to reduce the “buffet” effect, considering the characteristics of a low aspect ratio wing with thin airfoil and variable camber. 
   According to the present invention, on the contrary, a wing  18 ,  20  with a trapezoidal platform and mid aspect ratio (AR=4) is used, characterized by the presence of a saw tooth (shown as S in  FIG. 2 ) at 67.5% of the gross wing span. The modification, in comparison with standard wings, is first of all related to the leading edge radius, shown as R is  FIG. 6 , which was circular (prior art) and now becomes triangular in order to optimize the stagnation point position, in the presence of the leading edge  68  and of the “Leading Edge Droops”  66  deflected at mid angles of attack. 
   As clearly results from the configuration of  FIG. 6 , which shows an enlarged section along VI-VI line of  FIG. 2 , each wing  18 ,  20  presents a variable camber profile, both along the leading edge  66  (“Leading Edge Droop”) and along the trailing edge  70 , near the ailerons  56 . The ailerons  56  are scheduled only in transonic regime, to have a camber reduction in order to reduce compressibility effects. 
   Quantitatively, the design point of the chord percentage extension of the leading edge  68  is equal to 0.36% with a tolerance between +0.5% and −0.2% from the nominal value, while the design point of the gross wing span at which the chord extension is applied, in comparison with the standard solutions, is equal to 8.2%, with a tolerance between +10% and −5% from the nominal value. 
   Other characteristics of the aircraft  10  include the fuselage  12 , which in the afterbody includes the integration of the engine nozzles  60  and the presence of a fuselage tail, indicated by reference character  90  in  FIG. 3 . 
   The area near the engine nozzles  60  is also optimized in order to reduce the negative effects, in terms of drag and longitudinal stability/control, generated by the engine flow next to the fuselage side wall  12  and the horizontal tail  44 . 
   The aircraft  10  also features a tricycle landing gear, including a nose landing gear and a main landing gear. The nose landing gear is a leg strut, with four doors closing the bay, and backside retraction. The main landing gear has front side retraction in order to allow fuselage belly loads installation. 
   The aircraft  10  of the present invention fits an automatic flight control system (“Fly By Wire”), digital quadruplex redundant, which optimizes the performance and the flight qualities. The system improves the flight safety, limiting, automatically, flight regimes which could be uncomfortable to the pilot or could lead to lose control (“Carefree Handling”). From the above description, the characteristics of the aircraft configuration result in improved aerodynamic performance. 
   It is clear that other modifications can be applied to the structure of the subject aircraft, without exceeding the novelties included in the concept of the present invention. It is also clear that in applying the teachings of the subject invention, materials, shapes and dimensions of the above mentioned details could vary in compliance with the requirements and the same details could be replaced with other details having the same technical characteristics.