Abstract:
The invention provides an aircraft lifting surface with a torsion box ( 13 ) of a composite material comprising an upper skin ( 21 ), a lower skin ( 23 ), a front spar ( 18 ), a rear spar ( 20 ), one or more intermediate spars ( 19, 19′ ) and a plurality of transverse ribs ( 25, 25′, . . .  ) arranged between the rear spar ( 20 ) and its adjacent intermediate spar ( 19′ ) and/or between the front spar ( 18 ) and the adjacent intermediate spar ( 19 ) for improving its structural behavior. The invention also provides a manufacturing method of said torsion box.

Description:
FIELD OF THE INVENTION 
       [0001]    The present invention refers to a torsion box of an aircraft and more in particular to a torsion box of a lifting surface. 
       BACKGROUND OF THE INVENTION 
       [0002]    The structure of an aircraft lifting surface usually comprises a torsion box. 
         [0003]    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.). 
         [0004]    The torsion box is the primary structure responsible for supporting all loads involved (aerodynamic, fuel, dynamics, etc.) and comprises several structural elements. 
         [0005]    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 and other lifting surfaces can be manufactured using CFRP. 
         [0006]    The design of composite torsion boxes requires combining two perspectives of different nature: that of structural design and that of manufacture. 
         [0007]    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 join in the assembly plant following schemes similar to those used in the aeronautics industry when only metallic materials were used. 
         [0008]    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  and  1   b.    
         [0009]    The HTP is structured by leading edges  11 , torsion boxes  13  and trailing edges  15  with control surfaces (flaps, elevators, rudders, etc.). 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  16  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. 
         [0010]    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. 
         [0011]    Since analytical tools to obtain an optimal design of a torsion box of an aircraft tail plane made of composite materials taking into account all the variables involved and especially those related to their manufacture are not available at present, the aeronautics industry is constantly demanding new torsion box proposals and new manufacturing methods that improve efficiency and/or costs of known torsion boxes. 
         [0012]    The present invention is directed to the attention of that demand. 
       SUMMARY OF THE INVENTION 
       [0013]    It is an object of the present invention to provide a torsion box of a composite material for an aircraft lifting surface allowing weight and cost reductions with respect to known torsion boxes. 
         [0014]    It is another object of the present invention to provide a manufacturing method of said torsion box. 
         [0015]    In one aspect, these and another objects are met by a torsion box comprising an upper skin, a lower skin, a front spar, a rear spar, one or more intermediate spars and a plurality of transverse ribs arranged between the rear spar and its adjacent intermediate spar and/or between the front spar and the adjacent intermediate spar. The integration of said ribs in a multi-spar torsion box is a key feature of the invention. 
         [0016]    The upper and lower skins may include reinforcing stringers in all the cells delimited by spars without ribs. 
         [0017]    In the case of a torsion box of a tail plane, the transverse ribs placed between the rear spar and the adjacent intermediate spar are arranged to receive and distribute the loads generated by control configuration devices of the aircraft tail plane, to improve the torsional rigidity of the torsion box and to avoid great deformations of the torsion box. Similarly the transverse ribs arranged between the front spar and the adjacent intermediate spars are intended to improve the torsional rigidity of the torsion box and to avoid great deformations of the torsion box. 
         [0018]    This multi-spar and multi-rib configuration of the torsion box combines the manufacturing advantages of a multi-spar configuration with the structural advantages of a multi-rib configuration. 
         [0019]    In another aspect, the above-mentioned objects are met by a method of manufacturing said torsion box comprising the following steps: a) manufacturing separately a monolithic ensemble comprising all the structural elements of the torsion box with the exception of the rear spar and/or the front spar affected by said transverse ribs and said rear spar and/or front spar; b) joining said rear spar and/or front spar affected by said transverse ribs to the monolithic ensemble. Therefore much of the torsion box is manufactured in an integrated manner, reducing the amount of components and fasteners and consequently the torsion box weight and cost. 
         [0020]    In an embodiment the manufacturing method of said monolithic ensemble comprises the following steps: a) providing a set of laminated preforms of a composite material for forming said monolithic ensemble, 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 monolithic ensemble and a second set of tools for forming the open part of the monolithic ensemble and subjecting the curing assembly to an autoclave cycle to co-cure said laminated preforms; c) demoulding the first set of tools in a spanwise direction and the second set of tools in a chordwise direction. 
         [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. 2   a  is a schematic perspective view of a torsion box according to the present invention including ribs between the rear spar and the adjacent intermediate spar. 
           [0025]      FIG. 2   b  is a schematic plan view of the monolithic ensemble and of the rear spar that are manufactured separately and then joined according to the manufacturing method of this invention. 
           [0026]      FIG. 3   a  and  FIG. 4   a  are, respectively, schematic cross sections of an embodiment of the curing assembly of said monolithic ensemble by the planes A-A and B-B of  FIG. 2   b.    
           [0027]      FIGS. 3   b  and  4   b  are schematic cross sections of an embodiment of the monolithic ensemble obtained after the curing and the demoulding of the tooling by the planes A-A and B-B of  FIG. 2   b.    
           [0028]      FIGS. 5   a  and  5   b  are schematic cross sections of the tooling used to form laminated preforms having a C and a double C shape. 
           [0029]      FIG. 5   c  is a sketch of the process for obtaining a rib preform. 
           [0030]      FIG. 6   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,  FIG. 6   b  is a schematic perspective view of all these modules and  FIG. 6   c  is a schematic perspective view of the rib resulting from the integration of two rib preforms. 
           [0031]      FIG. 7   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. 7   b  is a schematic perspective view of all these modules. 
           [0032]      FIG. 8  is a schematic view of the demoulding process of the curing assembly in the case of having transversal ribs only close to the rear spar. 
           [0033]      FIGS. 9   a ,  9   b  and  9   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. 
           [0034]      FIGS. 10   a  and  11   a  are, respectively, schematic cross sections of two embodiments of the curing assembly of said monolithic ensemble by the plane A-A of  FIG. 2   b.    
           [0035]      FIGS. 10   b  and  11   b  are, respectively, schematic cross sections of two embodiments of the monolithic ensemble obtained after the curing and the demoulding of the tooling by the plane A-A of  FIG. 2   b.    
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0036]    In the following detailed description we would refer to the torsion box of an HTP but the invention is applicable to the torsion box of any lifting surface of an aircraft. 
         [0037]      FIG. 2   a  shows a composite torsion box  13  of an HTP according to an embodiment of the invention comprising the following structural elements:
       A front spar  18 , a rear spar  20  and intermediate spars  19 ,  19 ′.   An upper skin  21  and a lower skin  23 .   Several transverse ribs  25 ,  25 ′,  25 ″,  25 ′″between the rear spar  20  and its adjacent intermediate spar  19 ′.       
 
         [0041]    This configuration, which is very advantageous from a manufacturing standpoint, addresses the specific loading issues of the rear part of the torsion box which occur in many of the typical HTP architectures. 
         [0042]    In that sense, the transverse rib  25  is provided for receiving and distributing the loads from the pivot point of the rotation axis of the HTP, the ribs  25 ′,  25 ″ are provided for receiving and distributing the loads from the actuator devices of the HTP control surfaces and the rib  25 ′″ is provided to increase the torsional rigidity and to avoid great deformations of the torsion box  13 . Obviously the number and location of transverse ribs depends on the specific architecture of the HTP. 
         [0043]    The method for manufacturing the torsion box  13  according to the invention is based, firstly, on the separate manufacture of the rear spar  20  and of a monolithic ensemble  30  incorporating all the structural elements of the torsion box  13  except the rear spar  20  and, secondly, in their attachment by mechanical means such as, for example, rivets. 
         [0044]    The rear spar  20  is manufactured according to the method mentioned in the background section and the monolithic ensemble  30  by a method based on prepreg technology comprising the steps described below:
       Preparing the set of laminated preforms that will form the monolithic ensemble  30  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 element 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 in 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.       
 
         [0049]    For the embodiment of the monolithic ensemble  30  illustrated in  FIGS. 2   b ,  3   b  and  4   b , the laminated preforms used to manufacture it are the following:
       Laminated preforms  41 ,  43 ,  45 ,  47  having a double C-shaped transversal section to form the inner part of the monolithic ensemble  30  between the front spar  18  and the intermediate spar  19 ′ (see particularly  FIGS. 3   a ,  3   b ).   Laminated preforms  55 ,  57 ;  55 ′,  57 ′;  55 ″,  57 ″;  55 ′″,  57 ′″ having a C-shaped transversal section and a lateral wall in their inner end (see also  FIG. 6   b ) to form ribs  25 ,  25 ′,  25 ″,  25 ′″.   Laminated preforms  49 ,  49 ′,  49 ″,  49 ′″,  49 ″″, having a C-shaped transversal section to form, together with the ribs  25 ,  25 ′,  25 ″,  25 ′″, the inner part of the monolithic ensemble  30  between the intermediate spar  19 ′ and the rear end (see also  FIG. 6   b ). Alternatively, a single laminated preform  50  can be used (see  FIGS. 7   a  and  7   b ).   Laminated preforms  51 ,  53  with the shape of skins  21 ,  23  to form its outer part.       
 
         [0054]    The double C-shaped laminated preforms  41 ,  43 ,  45 ,  47  configured by a web, two primary flanges and two secondary flanges, are formed (see  FIG. 5   b ) bending the ends of a flat lay-up on a tooling  37  in two steps to get the primary flanges and the secondary flanges. The latter are those that form the reinforcing stringers  22 ,  24  of skins  21 ,  23 . 
         [0055]    The C-shaped laminated preforms  49 ,  49 ′,  49 ″,  49 ′″,  49 ″″ or the C-shaped laminated preform  50 , configured by a web and two flanges, are formed (see  FIG. 5   a ) bending the ends of a flat lay-up on a tooling  35  to get the flanges. 
         [0056]    The rib preforms  55 ,  57 ;  55 ′,  57 ′;  55 ″,  57 ″;  55 ′″,  57 ′″ configured by a web, two flanges and a lateral wall are formed bending a flat laminate.  FIG. 5   c  shows the bending operations—indicated by arrows F 1 , F 2 , F 3 —needed to form the flanges and the lateral wall of a rib preform  55  (the tooling is not shown).  FIG. 6   c  shows the rib  25  resulting from the integration of preforms  55 ,  57  which is configured by a web  27 , two flanges  28 ,  28 ′ and a lateral wall  29  having the same height than the web  27  and the same width than the flanges  28 ,  28 ′. 
         [0057]    The thickness and composite material of each laminated preform are defined according to the structural needs of the structural elements of the torsion box  13 . 
         [0058]    As illustrated in  FIGS. 3   a  and  4   a , said preforms are arranged on a tooling (see also  FIG. 8 ) forming a curing assembly  40  which will be subjected to an autoclave cycle to get the monolithic ensemble  30 . 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 ′.   Tools  65 ,  67 ,  69 ,  71 ,  73  extended on the spaces foreseen to be delimited by ribs  25 ,  25 ′,  25 ″,  25 ′″.  FIG. 6   a  shows particularly the assembly of the module comprising the rib preforms  57 ,  55 ′, the C-shaped preform  49 ′ and the tool  67 .       
 
         [0062]    As illustrated particularly in  FIG. 8 , tools  61 ,  63  are demoulded in the spanwise direction D 1  of the curing assembly  40  and tools  65 ,  67 ,  69 ,  71 ,  73  are demoulded in the chordwise direction D 2  of the curing assembly  40 . 
         [0063]    In the case of torsion boxes having skins  21 ,  23  with substantial curvature it may be desirable to divide the tools  65 ,  67 ,  69 ,  71 ,  73  into parts to facilitate the demoulding process. See  FIGS. 9   a ,  9   b ,  9   c  in which the tool  65  has been divided into three parts  65 ′,  65 ″,  65 ″ for demoulding the central part  65 ″ in the chordwise direction in the first place and the tools  65 ′,  65 ″ in the second place, separating them from the skins  21 ,  23  in a vertical direction in a first step and removing them in a chordwise direction in a second step. 
         [0064]      FIG. 10   b  shows another embodiment of a monolithic ensemble  30  according to the invention without stringers reinforcing the skins  21 ,  23  but with connecting flanges  36 ,  38  with the rear spar  20 . 
         [0065]    The laminated preforms used to manufacture it (see  FIG. 10   a ) are the following:
       Laminated preforms  42 ,  44 ,  46 ,  48  having a C-shaped transversal section to form the closed part of the monolithic ensemble  30 .   The same laminated preforms used in the previous embodiment to form the transverse ribs.   A single preform  54  having a double C-shaped transversal section to form, together with the ribs, the inside of the open part of the monolithic ensemble  30  or, alternatively, a set of preforms as in the modular configuration of the previous embodiment illustrated in  FIG. 6   b .   Laminated preforms  51 ,  53  with the shape of skins  21 ,  23  to form its outer part.       
 
         [0070]    The second set of tools comprises three tools  65 ′,  65 ″,  65 ″, . . . in each inner space of the open part of the curing assembly  40  to facilitate the chordwise demoulding. 
         [0071]      FIG. 11   b  shows another embodiment of a monolithic ensemble  30  according to the invention with stringers  22 ,  24  reinforcing the skins  21 ,  23  in the closed part of the monolithic ensemble  30  and connecting flanges  36 ,  38  with the rear spar  20 . 
         [0072]    The laminated preforms used to manufacture it (see  FIG. 11   a ) are the following:
       Laminated preforms  41 ,  43 ,  45 ,  47  having a double C-shaped transversal section to form the inner part of the monolithic ensemble  30  between the front spar  18  and the intermediate spar  19 ′.   The same laminated preforms used the previous embodiments to form the transverse ribs.       
 
         [0075]    A single preform  54  having a double C-shaped transversal section to form, together with the ribs, the inside of the open part of the monolithic ensemble  30  or, alternatively, a set of preforms as in the modular configuration of the previous embodiment illustrated in  FIG. 6   b.  
       Laminated preforms  51 ,  53  with the shape of skins  21 ,  23  to form its outer part.       
 
         [0077]    The second set of tools comprises three tools  65 ′,  65 ″,  65 ″, . . . in each inner space of the open part of the curing assembly  40  to facilitate their chordwise demoulding. 
         [0078]    After completing the demoulding process, the monolithic ensemble  30  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 doesn&#39;t have any defects. 
         [0079]    In other embodiments of the invention, the torsion box  13  may comprise transverse ribs between the front spar  18  and the middle spar  19 , alternatively or additionally to the transverse ribs between the rear spar  20  and the intermediate spar  19 ′, to reinforce the front part of the torsion box  13 . These embodiments will be manufactured following the guidelines of the manufacturing method described above. 
         [0080]    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.