Abstract:
The invention provides an aircraft lifting surface with a monolithic main supporting structure ( 14 ) of a composite material comprising an upper skin ( 21 ) including at least a part of the upper aerodynamic profile of the leading edge ( 11 ) and/or of the trailing edge ( 15 ), a lower skin, a front spar ( 18 ), a rear spar ( 20 ), a plurality of leading edge ribs and/or a plurality of trailing edge ribs. This main supporting structure allows a weight and cost reduction of aircraft lifting surfaces. The invention also provides a manufacturing method of said monolithic main supporting structure ( 14 ).

Description:
FIELD OF THE INVENTION 
       [0001]    The present invention refers to an aircraft lifting surface and more in particular to the main supporting structure of the lifting surface. 
       BACKGROUND OF THE INVENTION 
       [0002]    An aircraft lifting surface usually comprises a torsion box as its main supporting structure. For example, an aircraft tail plane (horizontal or vertical) is usually structured by a leading edge, a torsion box and a trailing edge with control surfaces (flaps, elevators, rudders, etc.). The torsion box is the main supporting structure responsible for supporting all loads involved (aerodynamic, fuel, dynamics, etc.) and comprises several structural elements. 
         [0003]    Composite materials with an organic matrix and continuous fibers, especially CFRP (Carbon Fiber Reinforced Plastic), are nowadays widely used in the aeronautical industry in a great variety of structural elements. Specifically, all the elements which make up the torsion boxes of aircraft tail planes can be manufactured using CFRP. 
         [0004]    The design of composite torsion boxes requires combining two perspectives of different nature: that of structural design and that of manufacture. 
         [0005]    The traditional approach is the design of the torsion box defining the structural elements that form it (skins, spars, stringers, ribs), the separate manufacture of these elements and their subsequent joint in the assembly plant following schemes similar to those used in the aeronautical industry when only metallic materials were used. 
         [0006]    The manufacture can be done using prepreg technology. In a first step, a flat lay-up of composite prepreg plies for each element is prepared. Then a laminated preform of the element with the required shape is obtained by means of a classical hot-forming process, being in some cases substituted by a press-forming process due to high curvatures. After getting the required shape, the laminated preform is cured in a male or female tooling depending on the tolerances required and the overall manufacturing cost. In the case of certain elements comprising sub-components cured separately, such as a rib and a vertical stiffener of it, a second curing cycle is needed for co-bonding said sub-components. Finally, after all the curing cycles, the element contours are trimmed getting the final geometry, and then the element is inspected by an ultrasonic system to assure its quality. The cost of a torsion box manufactured with said method is high because said steps shall be carried out independently for each structural element. Additionally, the cost related to the assembly of the torsion box is also high due to the long length and high complexity of the tasks required to install and to fit all structural elements together. This approach is being followed for manufacturing multi-rib torsion boxes such as that of the horizontal tail plane (HTP) shown in  FIGS. 1   a ,  1   b ,  1   c.    
         [0007]    The HTP is structured by leading edges  11 , torsion boxes  13  and trailing edges  15  with control surfaces (flaps, elevators, rudders, etc.). 
         [0008]    The leading edge is the structure responsible for keeping the aerodynamic surface with the torsion box surface, for supporting the static or cyclic structural loads involved and for protecting the torsion box from bird impacts. It is the part of the HTP surface that first contacts the air and the foremost edge of the airfoil. 
         [0009]    A known leading edge  11  comprises, on the one side, several ribs  10 , called leading edge ribs, attached to the front spar  18  of the torsion box  13  and, on the other side, an aerodynamic profile  12 —commonly known as “nose”-attached to the leading edge ribs  10  and to the flanges of the front spar  18  in order to keep the overall aerodynamic shape of the HTP. 
         [0010]    Similarly the trailing edge  15  comprises, on the one side, several ribs, called trailing edge ribs attached to the rear spar  20  of the torsion box  13  and, on the other side, an aerodynamic profile  16  attached to the trailing edge ribs and to the flanges of the rear spar  20  in order to keep the overall aerodynamic shape of the HTP between the torsion box and the control surfaces. 
         [0011]    The structural elements of torsion boxes  13  are upper and lower skins  21 ,  23  stiffened by longitudinal stringers, a front spar  18 , a rear spar  20  and transverse ribs  17  attached to the front and rear spars  18 ,  20  and to the upper and lower skins  21 ,  23  in order to keep the torsion box shape and reinforce the load introductions areas linked to the HTP structural arrangement in the aircraft and to the actuators for handling the HTP control surfaces. 
         [0012]    An alternative approach is to manufacture the whole or a part of a torsion box in an integrated manner for obtaining a monolithic ensemble comprising all or part of the structural elements of the torsion box. In this respect one example is described in WO 2008/132251 for a multi-spar torsion box. 
         [0013]    Due to the complexity of aircraft tail planes the aeronautics industry is constantly demanding new proposals and new manufacturing methods that improve efficiency and/or costs of known aircraft tail planes. 
         [0014]    The present invention is directed to the attention of that demand. 
       SUMMARY OF THE INVENTION 
       [0015]    It is an object of the present invention to provide a main supporting structure of an aircraft lifting surface of a composite material allowing weight and cost reductions with respect to a comparable structure of known aircraft lifting surfaces. 
         [0016]    It is another object of the present invention to provide a method of manufacturing said main supporting structure. 
         [0017]    In one aspect, these and another objects are met by a main supporting structure comprising an upper skin, a lower skin, a front spar, a rear spar (and optionally intermediate spars) and a plurality of leading and/or trailing edge ribs; the upper skin including at least a part of the upper aerodynamic profile of the leading edge and/or of the trailing edge; the main supporting structure being a monolithic structure (i.e. without joints). The spars ensure torsional stiffness and overall stability to withstand the required loads and the ribs keep the aerodynamic shape and support movable surfaces (if any). 
         [0018]    In embodiments of the invention, the lower skin includes at least a part of the lower aerodynamic profile of the leading edge and/or of the trailing edge. 
         [0019]    In embodiments of the invention, the upper and lower skins of the main supporting structure include reinforcing stringers in all the cells delimited by spars. 
         [0020]    In another aspect, the above-mentioned objects are met by a method of manufacturing said main supporting structure comprising the following steps: a) providing a set of laminated preforms of a composite material for forming said main supporting structure, each laminated preform being configured to form a part of it; b) arranging said laminated preforms in a curing assembly comprising a first set of tools for forming the closed part of the main supporting structure and a second set of tools for forming the open part of the main supporting structure; c) subjecting the curing assembly to an autoclave cycle to co-cure said laminated preforms; d) demoulding the first set of tools in a spanwise direction and the second set of tools in a chordwise direction. The invention therefore provides a high integrated solution to include leading and/or trailing edge ribs and leading and/or trailing edge aerodynamic profiles in a “one-shot” manufacturing process of a main supporting structure of an aircraft lifting surface of composite material, allowing the reduction of the amount of components and fasteners and consequently the weight and cost. 
         [0021]    Other desirable features and advantages of the invention will become apparent from the subsequent detailed description of the invention and the appended claims, in relation with the enclosed drawings. 
     
    
     
       DESCRIPTION OF THE FIGURES 
         [0022]      FIG. 1   a  is a perspective view of a known horizontal tail plane showing the torsion boxes, the leading edges and the trailing edges with control surfaces. 
           [0023]      FIG. 1   b  is a perspective view of a known torsion box, where the upper skin has been moved upwards to improve the visibility inside the box. 
           [0024]      FIG. 1   c  is perspective view of one side of the horizontal tail plane surface of  FIG. 1   a  with cutaways to improve the visibility of the leading edge structure showing the leading edge ribs and the leading edge profiles. 
           [0025]      FIG. 2  is a schematic perspective view of an embodiment of a main supporting structure according to the present invention comprising first and second leading edge ribs and first and second trailing edge ribs. 
           [0026]      FIG. 3   a  is a schematic perspective view of an embodiment of a main supporting structure according to the present invention comprising second trailing edge ribs. 
           [0027]      FIG. 3   b  is a schematic cross section of  FIG. 2   a  by plan C-C. 
           [0028]      FIGS. 4   a  and  5   a  are schematic cross sections of an embodiment of the curing assembly of the main supporting structure of  FIG. 3   a  by, respectively, the planes A-A and B-B. 
           [0029]      FIGS. 4   b  and  5   b  are schematic cross sections of the monolithic main supporting structure obtained after the curing and the demoulding of the tooling of the curing assembly by, respectively, the planes A-A and B-B of  FIG. 3   a.    
           [0030]      FIGS. 6   a  and  6   b  are, respectively, schematic cross sections of another embodiment of the curing assembly of said main supporting structure and of the monolithic main supporting structure obtained after the curing and the demoulding of the tooling of the curing assembly by, respectively, the planes A-A and B-B of  FIG. 3   a.    
           [0031]      FIGS. 7   a  and  7   b  are schematic cross sections of the tooling used to form laminated preforms having a C and a double C shape. 
           [0032]      FIG. 7   c  is a sketch of the process for obtaining a rib laminated preform. 
           [0033]      FIG. 8   a  is a diagram illustrating the arrangement of the preforms of one of the modules to be integrated in the rear part of the main supporting structure, 
           [0034]      FIG. 8   b  is a schematic perspective view of the set of said modules (assuming that they have the same dimensions) and  FIG. 8   c  is a schematic perspective view of the rib resulting from the integration of two rib laminated preforms. 
           [0035]      FIGS. 9   a  is a diagram illustrating the arrangement of the preforms of one of the modules to be integrated in the rear part of the torsion box in another embodiment of the invention and  FIG. 9   b  is a schematic perspective view of all these modules. 
           [0036]      FIG. 10  is a schematic view of the demoulding process of the curing assembly. 
           [0037]      FIGS. 11   a ,  11   b  and  11   c  are schematic representations of the demoulding process of the tooling of the open part of the monolithic ensemble in a particular embodiment of said tooling. 
           [0038]      FIG. 12  is a schematic cross section of an embodiment of a main supporting structure according to the invention having trailing edge ribs which are covered by its upper skin and are not covered by its lower skin. 
           [0039]      FIG. 13  is a schematic cross section of a main supporting structure according to the invention having leading and trailing edge ribs which are covered by its upper skin and are not covered by its lower skin. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0040]    In the following detailed description we would refer to the main supporting structure of an HTP but the invention is applicable to the main supporting structure of any lifting surface of an aircraft. 
         [0041]      FIG. 2  shows a monolithic main supporting structure  14  of an HTP according to an embodiment of the invention comprising the following structural elements:
       A front spar  18  and a rear spar  20 .   An upper skin  21  and a lower skin  23  including a part of the aerodynamic profiles of the leading edge  11  and the trailing edge  15 .   First leading edge ribs  22  extended inside the leading edge  11  and second leading edge ribs  24  extended inside a region of the leading edge  11  covered by the upper skin  21  and the lower skin  23 .   First trailing edge ribs  25  extended inside the trailing edge  15  and second trailing edge ribs  26  extended inside a region of the trailing edge  15  covered by the upper skin  21  and the lower skin  23 .       
 
         [0046]    Consequently the main supporting structure  14  comprises the torsion box of known HTP plus part of the leading edge and of the trailing edge. 
         [0047]    This configuration, which is very advantageous from a manufacturing standpoint, addresses the specific loading issues of the front and rear parts of the torsion box which occur in many of the typical HTP architectures. Obviously the number and location of leading and trailing edge ribs depends on the specific architecture of the HTP. 
         [0048]    Other embodiments of a monolithic main supporting structure  14  of an HTP according to the invention comprise different configurations of its front and rear sides including or not including all or part of the above-mentioned leading and trailing edge ribs, and including or not including parts of the aerodynamic profile of the leading edge  11  and/or of the trailing edge  15 . One of them is shown in  FIGS. 3   a  and  3   b  and comprises the following structural elements:
       A front spar  18 , a rear spar  20  and intermediate spars  19 ,  19 ′.   Several trailing edge ribs  26  including both structural ribs and bearing ribs (for example the ribs which support the elevator hinge line).   An upper skin  21  and a lower skin  23  including a part of the aerodynamic profile of the trailing edge  15  covering the trailing edge ribs  26 .       
 
         [0052]    Other embodiments of the main supporting structure  14  with different configurations of the upper and lower skin are shown in  FIGS. 12 and 13 . 
         [0053]      FIG. 12  show an embodiment where only the upper skin  21  covers the trailing edge ribs  26 . 
         [0054]      FIG. 13  shows an embodiment having leading edge ribs  22  and trailing edge ribs  26  where only the upper skin  21  covers the trailing edge ribs  26  and part of the leading edge ribs  22 . 
         [0055]    A method for manufacturing the monolithic main supporting structure  14  shown in  FIGS. 3   a  and  3   b  according to the invention is based on prepreg technology and comprises the following steps:
       Preparing a set of laminated preforms that will form the monolithic main supporting structure  14 , laying-up for each of them a flat lay-up of composite prepreg plies and subjecting the flat lay-up to a hot-forming process on a suitable tool to give it the desired shape or performing the desired lay-up over a surface with the desired shape. The term “laminated preform” as used in this specification designates a composite that is intended to be integrated with other elements in the manufacturing process of the product to which it belongs.   Arranging together all the laminated preforms on a curing assembly  40  with a suitable tooling and subjecting the curing assembly  40  to an autoclave cycle to co-cure the laminated preforms.   Demoulding the tooling.   Trimming and inspecting the assembly.       
 
         [0060]    The laminated preforms used to manufacture the monolithic main supporting structure  14  of  FIGS. 4   b  and  5   b , comprising upper and lower skins  21 ,  23 , with reinforcing stringers  32 ,  34  in all the closed cells, are the following (see  FIGS. 4   a ,  5   a ):
       Laminated preforms  41 ,  43 ,  45 ,  47 ,  49 ,  51  having a double C-shaped transversal section to form the inner part of the monolithic main supporting structure  14  between the front spar  18  and the rear spar  20 .   Laminated preforms  53  having a C-shaped transversal section to form the inner part of the monolithic main supporting structure  14  between the rear spar  20  and the rear end together with pairs of laminated preforms  35 ,  37  having a C-shaped transversal section and a lateral wall in their inner ends to form the trailing edge ribs  26  (see also  FIGS. 8   a ,  8   b  and  8   c ). In the embodiment shown in  FIGS. 9   a  and  9   b  a single laminate preform  54  having a C-shaped transversal section is used instead of said laminated preforms  53 .   Laminated preforms  57 ,  59  with the shape of upper and lower skins  21 ,  23  to form its outer part.       
 
         [0064]      FIG. 6   b  shows another embodiment of the monolithic main supporting structure  14  comprising upper and lower skins  21 ,  23  without reinforcing stringers.  FIG. 6   a  shows the set of laminated preforms for this embodiment comprising laminated preforms  42 ,  44 ,  46 ,  48 ,  50 ,  52  having a C-shaped transversal section instead of the laminated preforms  41 ,  43 ,  45 ,  47 ,  49 ,  51  of the embodiment shown in  FIG. 4   a.    
         [0065]    The double C-shaped laminated preforms  41 ,  43 ,  45 ,  47 ,  49 ,  51 , configured by a web, two primary flanges and two secondary flanges, are formed (see  FIG. 7   b ) bending the ends of a flat lay-up on a tooling  39  in two steps to get the primary flanges and the secondary flanges. The latter are those that form the reinforcing stringers  32 ,  34  of upper and lower skins  21 ,  23  (see  FIG. 4   b ). 
         [0066]    The C-shaped laminated preforms  53 ,  54 ,  42 ,  44 ,  46 ,  48 ,  50 ,  52  configured by a web and two flanges, are formed (see  FIG. 7   a ) bending the ends of a flat lay-up on a tooling  38  to get the flanges. 
         [0067]    The rib preforms  35 ,  37  configured by a web, two flanges and a lateral wall are formed bending a flat laminate.  FIG. 7   c  shows the bending operations -indicated by arrows F1, F2, F3—needed to form the flanges and the lateral wall of a rib laminated preform  35  (the tooling is not shown). 
         [0068]      FIG. 8   c  shows the rib  26  resulting from the integration of preforms  35 ,  37  which is configured by a web  27 , two flanges  28 ,  28 ′ and a lateral wall  29  having the same height as the web  27  and the same width as the flanges  28 ,  28 ′. 
         [0069]    The thickness and composite material of each laminated preform are defined according to the structural needs of the structural elements of the main supporting structure  14 . 
         [0070]    As illustrated in  FIGS. 4   a ,  5   a  and  6   a  said preforms are arranged on a tooling (see also  FIG. 10 ) forming a curing assembly  40  which will be subjected to an autoclave cycle to get the main supporting structure  14 . 
         [0071]    Said tooling comprises the following elements:
       A tool  61  extended on the space foreseen to be delimited by the front spar  18  and the intermediate spar  19 .   A tool  63  extended on the space foreseen to be delimited by the intermediate spars  19 ,  19 ′.   A tool  65  extended along the space foreseen to be delimited between the intermediate spar  19 ′ and the rear spar  20 .   Tools  67 ,  69 ,  71 ,  73 ,  75 ,  77  extended on the spaces foreseen to be delimited by ribs  26 .  FIG. 8   a  shows particularly the assembly of the module corresponding to the tool  69  with the rib preforms  37 ,  35  and the C-shaped preform  53 .       
 
         [0076]    As illustrated particularly in  FIG. 10 , tools  61 ,  63 ,  65  are demoulded in the spanwise direction D1 of the curing assembly  40  and tools  67 ,  69 ,  71 ,  73 ,  75 ,  77  are demoulded in the chordwise direction D2 of the curing assembly  40 . 
         [0077]    In the case of a main supporting structure  14  having upper and lower skins  21 ,  23  with substantial curvature may be desirable to divide the tools  67 ,  69 ,  71 ,  73 ,  75 ,  77  into parts to facilitate the demoulding process. See  FIGS. 11   a ,  11   b ,  11   c  in which the tool  67  has been divided into three parts  67 ′,  67 ″,  67 ′″ for demoulding the central part  67 ″ in the chordwise direction in the first place and the tools  67 ′,  67 ′″ in the second place, separating them from the upper and lower skins  21 ,  23  in a vertical direction in a first step and removing them in a chordwise direction in a second step. 
         [0078]    In another embodiment of the invention for a main supporting structure  14  having upper and skins  21 ,  23  with substantial curvature, the part of the lower skin  23  covering the trailing edge ribs  26  is joined to the rest of the lower skin  23  in an articulated manner (for example by means of hinges) so that the tools  67 ,  69 ,  71 ,  73 ,  75 ,  77  can be demoulded in a vertical direction. 
         [0079]    After completing the demoulding process, the monolithic main supporting structure  14  is located in the trimming machine in order to get the final geometry and is subjected to an automatic ultrasonic inspection for verifying that it does has not have any defects. 
         [0080]    These manufacturing methods are applicable mutatis mutandi to other embodiments of the main supporting structure according to the invention. 
         [0081]    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.