Abstract:
An aircraft lifting surface attached to the rear or frontal end of the aircraft fuselage with a variable sweep angle α in an inboard part and with a constant sweep angle α 1  in an outboard part. The aircraft lifting surface can be for example a horizontal tail plane or a vertical tail plane attached to the rear end fuselage or a canard attached to the frontal end fuselage.

Description:
CROSS-REFERENCES TO RELATED APPLICATIONS 
       [0001]    This application claims the benefit of the European patent application No. 12382282.7 filed on Jul. 16, 2012, the entire disclosures of which are incorporated herein by way of reference. 
       FIELD OF THE INVENTION 
       [0002]    The present invention relates to aircraft lifting surfaces and more in particular to aircraft with a lifting surface attached to an end of the fuselage such as a horizontal tail plane, a vertical tail plane or a canard. 
       BACKGROUND OF THE INVENTION 
       [0003]    The performance of the horizontal tail plane (HTP), the vertical tail plane (VTP) and other lifting surfaces attached to the ends of aircraft fuselages is one of the more important issues in global aircraft design because said surfaces are used as control and stabilizing surfaces that must provide control and stabilizing forces in the complete flight domain. 
         [0004]    The aim of a good aerodynamic design for a lifting surface is to control the interferences with the fuselage which are sources of aerodynamic drag and loss of lift. In this respect, the sweep angle of a lifting surface is a key feature of its design. 
         [0005]    Aircraft configurations with forward swept and backward swept horizontal tail planes at different sweep angles are known in the art. In relation to commercial aircraft configured with a tubular fuselage, a wing, an empennage with HTP and VTP, such as the A320 or the A380, all known configurations include HTP/VTP with a constant sweep angle distribution along the span. 
         [0006]    In these configurations and due to the interference with the fuselage, the inner sections of the HTP/VTP are not working at the same flow conditions than the outer sections, providing room for further optimization. This effect is more pronounced the greater the change in the area of the cross section of the rear fuselage along the zone to which the HTP/VTP is attached. 
       SUMMARY OF THE INVENTION 
       [0007]    It is an object of the present invention to provide a lifting surface attached to an end of the fuselage of an aircraft optimized in size. 
         [0008]    It is another object of the present invention to provide a lifting surface attached to an end of the fuselage of an aircraft having an improved lift curve slope with respect to known lifting surfaces. 
         [0009]    These and other objects are met by a lifting surface attached to the frontal end or to the rear end (that have a variable cross-sectional area) of a tubular-shaped fuselage of an aircraft which is configured with a variable sweep angle α in an inboard part and with a constant sweep angle α 1  in an outboard part. 
         [0010]    The lifting surface can be either a backward-swept lifting surface, as happens in the majority of commercial aircraft, or a forward-swept lifting surface. 
         [0011]    In an embodiment for a lifting surface attached to the fuselage rear end (such as a HTP or a VTP), the sweep angle α in the inboard part (which is variable along its span) is lower than the constant sweep angle α 1  in the outboard part. The local Mach number distribution along the span of the lifting surface due to the interference with the fuselage (lower Mach numbers in the inboard part than in the outboard part) allows a reduction of the sweep angle in the inboard part that increases the lift curve slope of the lifting surface. 
         [0012]    Advantageously, the variable sweep angle α in the inboard part of the lifting surface increases along its span. A progressive increment of the sweep angle α in the inboard part up to the constant value α 1  in the outboard part provides an optimized design of the lifting surface. 
         [0013]    Advantageously, the lifting surface comprises a leading edge, a torsion box and a trailing edge and the torsion box comprises straight frontal and rear spars. The variable sweep angle in the inboard part is thus compatible with a torsion box comprising straight spars. 
         [0014]    In an embodiment for a lifting surface attached to the fuselage frontal end (such as a canard) the sweep angle α in the inboard part (which is variable along its span) is greater than the constant sweep angle α 1  in the outboard part. The increase of the sweep angle in the inboard part decreases the lift curve slope but allows delaying the adverse effects of compressibility and decreasing the sweep angle of the outboard part. 
         [0015]    Advantageously the sweep angle in the inboard part of the lifting surface attached to the frontal end decreases along its span. A progressive decrement of the sweep angle α in the inboard part up to the constant value α 1  in the outboard part provides an optimized design of the lifting surface. 
         [0016]    Other desirable features and advantages of the aircraft according to this invention will become apparent from the subsequent detailed description of the invention and the appended claims, in relation with the enclosed drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0017]      FIG. 1  is a schematic plan view of half of the fuselage rear end of a known aircraft with a backward-swept horizontal tail plane. 
           [0018]      FIG. 2  is a Mach number vs. Span diagram for the horizontal tail plane of  FIG. 1 . 
           [0019]      FIG. 3  is a schematic plan view of half of the fuselage rear end of an aircraft with a backward-swept horizontal tail plane according to the present invention. 
           [0020]      FIG. 4  is an enlarged view of the projection on a horizontal plane of a line at the 25% of the chord in the inboard part of the horizontal tail plane according to the present invention. 
           [0021]      FIG. 5  a schematic plan view of half of the fuselage frontal end of an aircraft with a backward-swept canard. 
           [0022]      FIG. 6  is an enlarged view of the projection on a horizontal plane of a line at the 25% of the chord in the inboard part of a canard according to the present invention. 
       
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
       [0023]    A detailed description of the invention for a backward swept HTP follows. 
         [0024]      FIG. 1  illustrates a known aircraft  9  with a HTP  13  attached to the fuselage rear end  11 . 
         [0025]    The HTP  13  comprises a leading edge  21 , a torsion box  25 , and a trailing edge  23 . The torsion box  25  comprise a frontal spar  31 , a rear spar  33 , ribs  35  and upper and lower skins stiffened by stringers (not shown). The upper and lower skins are joined to the leading edge  21  and to the trailing edge  23  forming the aerodynamic contour of the HTP  13 . 
         [0026]    The HTP  13  is configured with a constant backward sweep angle α 1 , i.e. with a constant sweep angle α 1  greater than 90°. The sweep angle is the angle formed between the aircraft plane of symmetry  19  and the projection line  17  of a reference line of points located at 25% of the local chord of the HTP  13  on a plane perpendicular to the aircraft plane of symmetry  19 . 
         [0027]    The sweep angle of aircraft airfoils is a design feature of aircraft that fly at speeds approaching the speed of sound, and it is motivated by aerodynamic considerations. The aerodynamic advantage of a backward sweep angle is that the adverse effects of compressibility, caused by the over speed of the flow over the aerodynamic profile, which grow as the relative thickness of that profile increases, are mainly dependent on the component of the airflow velocity that is essentially perpendicular to the line of 25% of the chord line of the aircraft airfoil. This velocity component decreases as the sweep angle increases (in absolute value, either positively for backward sweep or negatively for forward sweep). 
         [0028]    Therefore, for a given flight speed, an airfoil with a given sweep angle will be subjected to lower compressibility effects. This effect allows the use of a bigger relative profile thickness, defined as the ratio between the maximum thickness of the profile and its length in the flight or chord direction, resulting in a lower structural weight of the airfoil because of a better structural efficiency. However, in the flight at high speed that is characteristic of large modern commercial aircraft, airfoils with large relative thicknesses of the aerodynamic profiles magnify the adverse effects of air compressibility, which can be manifested as shock waves on the airfoil, with an associated increase of the aerodynamic drag, loss of control capability and other adverse flight phenomena. Therefore, the backward or forward sweep angle of airfoils serves to achieve a design balance between their structural weight and acceptable in-flight performance at speeds approaching the speed of sound. 
         [0029]    However, analysing the performance of the known backward-swept HTP  13  it has been noted that in some cases the distribution of the Mach number M along the span S follows the curve  40  shown in  FIG. 2 . The inboard sections of the HTP  13  are thus working at lower Mach numbers than the rest of the sections; therefore those sections do not need the sweep angle values of the outboard sections as the compressibility effects are naturally delayed by the interaction with the fuselage recompression. As a side effect this causes the lift curve slope of the HTP  13  to be reduced with respect to the one that hypothetically would be obtained if the local Mach number were constant across the span and equal to the flight Mach number. 
         [0030]    It is believed that this behaviour is due to the interference of the airflow with the rear end fuselage  11  because of its curved shape (in a plan view) and the recompression which occurs as the flow approach to the fuselage end. The effect is more pronounced the greater the change in the area of the cross section along the zone of the rear fuselage to which the HTP is attached. 
         [0031]    The opposite behaviour can be found in lifting surfaces attached at the frontal end of the fuselage as the flow expands from the forward stagnation point onwards. However it usually does not happen in a wing because it is attached to a cylindrical-shaped fuselage where no expansion or recompression occurs due to the fuselage shape. Of course the cylindrical fuselage affects the flow on the wing but the effect is usually more related to the change of the flow direction than to the change of the local Mach number which typically occurs at the rear and at the forward end of the fuselage. 
         [0032]      FIG. 3  illustrates an aircraft  9  with an HTP  43  attached to the fuselage rear end  11  configured according to this invention. 
         [0033]    The HTP  43  comprises a leading edge  51 , a torsion box  55 , and a trailing edge  53 . 
         [0034]    The HTP  43  is configured with an inboard section  45  having an increasing sweep angle α along the span and an outboard section  47  having a constant sweep angle α 1 , the constant angle α 1  being greater than any value of the sweep angle α in the inboard section  45  (see  FIG. 4 ). 
         [0035]    The reduced sweep angles in the inboard section  45  with respect to the sweep angle of the outboard section  47  increases the lift curve slope of the HTP  43  with respect to the HTP  13  of the prior art, consequently allowing a size reduction in case the size of the HTP  13  would be a relevant design variable. 
         [0036]    In the embodiment shown in  FIG. 3 , the torsion box  55 , comprising a frontal spar  61 , a rear spar  63 , ribs  65  and upper and lower skins stiffened by stringers (not shown), has the same configuration as the torsion box  23  of the HTP  13  of the prior art because the variation of the sweep angle in the inboard section  45  with respect to the HTP  13  does not require a modification in the configuration of the torsion box. In other embodiments the torsion box  55  may have a different configuration. 
         [0037]    The above description of the invention for a backward/forward swept HTP is also applicable mutatis mutandi to a forward HTP and also to a backward/forward sweep VTP. 
         [0038]      FIGS. 5 and 6  show a backward swept canard  73  attached to a fuselage frontal end  10  whose cross-sectional area increases continuously along its length. 
         [0039]    The canard  73  is configured with an inboard section  75  having a decreasing sweep angle α along the span and an outboard section  77  having a constant sweep angle α 1 , the constant angle α 1  being lower than any value of the sweep angle α in the inboard section  75 . 
         [0040]    The sweep angle is the angle formed between the aircraft plane of symmetry  19  and the projection line  70  of a reference line of points located at 25% of the local chord of the canard  73  on a plane perpendicular to the aircraft plane of symmetry  19 . 
         [0041]    The increment of the sweep angle in the inboard section  75  with respect to the sweep angle of the outboard section  77  allows delaying the compressibility effects, magnified by the local expansion of the nose fuselage. The lift curve slope of the canard  73  can be increased by reducing the sweep angle of the outboard sections from the values which are needed inboard, allowing the reduction of the size of the canard  73  if its size is a relevant design variable. 
         [0042]    Although the present invention has been described in connection with various embodiments, it will be appreciated from the specification that various combinations of elements, variations or improvements therein may be made, and are within the scope of the invention.