Abstract:
An aircraft structure including stiffened panels assembled at the junction, without continuity of the stiffeners. Multiple panels are used to form the fuselage, tail units and wings of an aircraft. These panels include stiffeners which are interrupted at each panel junction. However, the stresses experienced by the stiffeners must be transmitted despite these interruptions. Known solutions require the use of at least one additional part per stiffener and per interruption. In order to solve this problem, the disclosed embodiments include a doubler necessitating at most one additional part per panel junction area for all of the interrupted stiffeners in this area.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is the National Stage of International Application No. PCT/FR2008/051631 International Filing Date, 12 Sep. 2008, which designated the United States of American, and which International Application was published under PCT Article 21 (s) as WO Publication No. WO2009/050358 A2 and which claims priority from, and the benefit of, French Application No. 200758423 filed on 18 Oct. 2007, the disclosures of which are incorporated herein by reference in their entireties. 
     BACKGROUND 
     The aspects of the disclosed embodiments relate to the field of structures comprising stiffened panels. This relates to panel essentially comprised of a thin coating, whose structural stability is ensured by elements relating to the coating. 
     More particularly, the disclosed embodiments relate to structures in which the stiffened panels, such as those used for producing the aircraft fuselage, are assembled at the junction of said panels without stiffener continuity. 
     For explanatory purposes regarding the state of art relating to the assembly of stiffened panels, such as, for example, one embodiment of the disclosed embodiments, the case of the assembly of stiffened panels aimed at producing an aircraft fuselage shall be described. 
     An aircraft fuselage is typically found in so-called “hull” structures, in particular for reasons related to lighter masses, which are essential in the field of aeronautics. On conventional aircraft, the fuselage comprises one substantially constant section on one part of its length, which gives the fuselage an overall well-known cylindrical shape. 
     For industrial and logistic reasons, such a fuselage is often produced from multiple cylindrical sections or from multiple section panels according to the type of assembly provided for by the industrial process. When the fuselage is produced from multiple sections, each section can itself be comprised of multiple panels. 
     In order to produce a rigid but light-weight structure, the hull structures generally comprise a relatively thin coating located on the inner wall of the cylinder. This coating is also referred to as skin. The hull structures also comprise structural elements attached to said coating, aimed at ensuring both the resistance and stability of said structures. In one aircraft fuselage structure, the structural elements substantially directed in the direction of the fuselage cylinder generators are referred to as stiffeners. The structural elements located on a substantially normal plane to said generators are referred to as frames. 
     With aircraft and for aerodynamic reasons, the stiffeners and frames and generally located inside the fuselage, and are therefore attached to the coating, itself located on the inner wall of the sections of fuselage. 
     When the panels are assembled to form a section or a fuselage, these panels are generally already equipped with stiffeners. The stiffeners therefore stop at the edges of the panel located on the same side as the stiffener ends, said panel edges being referred to, by extension, as panel ends. These panels comprised of a coating and stiffeners, are referred to as self-stiffened panels. 
     During the assembly operation for the sections or panels in order to produce a fuselage, the stresses in the coating and stiffeners must be transmitted from one section to the other or from one panel to the other. This refers in particular to tensile, compressive and/or shearing stresses. 
     In order to ensure the transfer of these stresses, one solution consists in creating coating and stiffener continuity at the level of a panel junction. This continuity is obtained by means of junction parts, on the one hand from panel to panel, and on the other hand from stiffener to stiffener. Said junction parts are produced in such a way as to preserve a transversal cross-section and constant inertia over the entire length of the structure. 
     One method generally used consists in attaching a plate to the coating of the two panels, said plate taking on the curve of the panels and partially covering the two panels positioned end to end. When the fuselage is assembled in sections, each plate, referred to as shroud, covers all or part of the fuselage perimeter at the level of the junction. 
     In order to create a connection between two stiffeners located opposite each other at the level of the junction, a specific part, referred to as a batten, is attached between the two stiffeners. A batten has a cross-section generally similar to that of the stiffeners concerned by the junction and covers each stiffener over a long enough distance to efficiently transmit the stresses from one stiffener to the other. 
     One of the difficulties connected to this type of junction is due to the poor alignment of the stiffeners that must be assembled at the level of a junction. 
     The unavoidable provisions regarding the dimensions of the stiffeners and their positions on the coatings, connected to the panel manufacturing and assembly methods, do not guarantee the precise alignment of the stiffeners between the two panels or the two sections being assembled. 
     One known solution consists in not attaching the stiffeners to the coatings over their entire length. A long enough length is left free at the ends of said stiffeners that must be battened. After having positioned the panels to be assembled, the stiffeners can thus be distorted within the limits of their field of elasticity, in order to align them before completing their assembly to the panels and batten. 
     This solution therefore requires particular assembly operations and cannot be performed in situations where the stiffeners are attached over their entire length, as, for example, with welded or bonded stiffeners, in particular in the case of structures made out of composite materials. 
     In these cases, the panels must be produced with very strict dimensional tolerances. This solution is limited to large-scale dimensions and always turns out to be very expensive. The alignment faults can also be corrected with blocks. The implementation of said blocks is a delicate and long procedure, requiring, when using a polymerisable mastic, a waiting time detrimental to the duration of the assemblies. 
     In addition, some stiffener shapes, particularly used in structures made out of composite materials, have closed cross-sections. Once assembled with a panel, the inside of the stiffener can no longer be accessed. This is the general case for stiffeners comprising two sole plates and a body connecting said two sole plates. These stiffeners are referred to as omega-shaped stiffeners due to their characteristic cross-section with a shape similar to a capital omega (Ω). 
     Without any possibility of inspecting the inside of such stiffeners, attachments are not recommended for unblocking the inside of these stiffeners. 
     Moreover, in some instances, the junction is produced in the presence of a reinforcing frame. In these cases, in order to ensure the passage of battens from one end of the stiffener to the other, openings must be made through the reinforcing frame. These openings reduce the level of structural resistance of the reinforcing frame. In addition, the presence of the frame significantly increases the complexity of the batten assembly operations and that of the section assembly operations. 
     SUMMARY 
     The disclosed embodiments offer a solution to resolving these difficulties in the prior art. The purpose of the disclosed embodiments is therefore to enable structures such as an aircraft fuselage to be simply and quickly manufactured and installed with these interrupted stiffeners. Another purpose of the disclosed embodiments is to enable the stresses to be efficiently transmitted, despite the interruptions in the stiffeners. 
     Moreover, the purpose of the disclosed embodiments is to avoid blind attachments, i.e. to avoid attachments that open up the inside of stiffeners with closed cross-sections, such as omega-shaped stiffeners. Another purpose of the disclosed embodiments is to provide junctions that do not require openings to be created within the frame. 
     In order to resolve these problems, the disclosed embodiments provide for the presence of a doubler. The purpose of a doubler is to lower the stress of the stiffener towards the coating or skin of the panel. In order to enable the progressive transfer of the stresses, the stiffeners are progressively interrupted. Such a doubler thus unloads the stiffeners so that the stiffener is almost completely free from stress at the level of the interruption of said stiffener, and that the shroud is sufficient in transmitting the stress from one panel to the other. 
     The disclosed embodiments aim at reducing the “level of stiffening” via the progressive interruption of the stiffeners. According to the disclosed embodiments, the stiffeners are systematically widened in the junction area. This enables the misalignments to be absorbed and to guarantee the presence of enough material between the attachments and the edges of the parts for the efficient transmission of the stresses. 
     The disclosed embodiments have multiple modes of embodiment, in particular at the level of the doubler. The doubler can be an additional part bonded above the skin, battened above the skin, or integrated underneath the skin of the panel. According to another mode of embodiment, the role of the doubler can be performed by the presence of an additional level of thickness of the skin underneath the stiffeners, or the doubler can even be integrated into the shroud. In a similar manner, the stiffener sole plates can be widened in order to touch each other and perform the role of the doubler, or even in some cases, no doubler could be used with a right-hand or toothed shroud. 
     The disclosed embodiments therefore relate to an aircraft structure comprising a first stiffened panel, said first panel comprising a skin and at least one stiffener, the stiffener of the first panel comprising at least two sole plates attached to one side of the skin of said panel, the stiffener of the first panel extending according to a longitudinal axis of the first panel, the stiffener of the first panel being interrupted, at least one second stiffened panel, said second panel comprising a skin and at least one stiffener, the stiffener of the second panel comprising at least two sole plates attached to one side of the skin of said panel, the stiffener of the second panel extending according to a longitudinal axis of the second panel, the stiffener of the second panel being interrupted, in said aircraft structure, the first panel and the second panel are assembled in such a way that their ends are placed close to each other, thus forming a line of interface, the stiffener of the first panel being opposite the stiffener of the second panel, said stiffeners being substantially aligned according to a direction substantially parallel to the longitudinal axis of the panels, a shroud partially covering the first and second panels, this shroud being on the one hand assembled on the first panel and on the other hand assembled on the second panel, said aircraft structure being characterised in that the structure partially covers at least two sole plates of the stiffener of the first panel and at least two sole plates of the stiffener of the second panel, said structure comprising at least one doubler, said doubler extending along the panels over a distance, according to the longitudinal axes of said panels, at least equal to the distance covered by the shroud according to these longitudinal axes, on either side of the line of interface of said panels, the doubler associated to the stiffener sole plates forming a continuous and regular support surface for the shroud. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The disclosed embodiments will be better understood after reading the following description and after examining the accompanying figures. These are presented as a rough guide and in no way as a limited guide to the disclosed embodiments. The figures show: 
         FIG. 1 : A profile view of an aircraft with a zoomed image of the elements of the structure comprising the self-stiffened panels; 
         FIGS. 2A and 2B : A representation of two examples of transversal cross-sections of panels comprising the O-shaped stiffeners; 
         FIG. 3 : An overhead schematic view of the junction area of the fuselage panels according to a first mode of embodiment of the disclosed embodiments; 
         FIG. 4 : A cross-section between two side-by-side stiffeners on the foot of the doubler according to this first mode of embodiment of the disclosed embodiments; 
         FIG. 5 : A cross-section of the panel junction area between the two stiffeners according to this first mode of embodiment of the disclosed embodiments; 
         FIG. 6 : An overhead schematic view of the junction area of the fuselage panels according to a second mode of embodiment of the disclosed embodiments; 
         FIG. 7 : A cross-section between two stiffeners according to this second mode of embodiment of the disclosed embodiments; 
         FIG. 8 : An overhead schematic view of the junction area of the fuselage panels according to a third mode of embodiment of the disclosed embodiments; 
         FIG. 9 : A cross-section between two stiffeners according to this third mode of embodiment of the disclosed embodiments; 
         FIG. 10 : An overhead schematic view of the junction area of the fuselage panels according to a fourth mode of embodiment of the disclosed embodiments; 
         FIG. 11 : A cross-section between two stiffeners according to this fourth mode of embodiment of the disclosed embodiments; 
         FIG. 12 : An overhead schematic view of the junction area of the fuselage panels according to a fifth mode of embodiment of the disclosed embodiments; 
         FIG. 13 : A cross-section between two stiffeners according to this fifth mode of embodiment of the disclosed embodiments; 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  represents a profile view of an aircraft with a zoomed image of the elements of the structure comprising the self-stiffened panels. Aircraft  1  comprises a fuselage  2  generally with a structure referred to as a “hull”. The detailed description of the disclosed embodiments will, from this point onwards, be applied to the case of the structure of fuselage  2 , this case capable of being easily adapted by one of ordinary skill in the art to other structures comprising self-stiffened panels of an aircraft  1 . 
     A fuselage  2  is produced from assembled sections  4  or panels  5 . More particularly, these sections  4  or panels  5  are connected between themselves to form the main structure of the fuselage  2 . 
     The wings or even vertical or horizontal tail units can also be produced from panels  5  in the case of an aircraft  1 . Moreover, a section  4  can be produced from panels  5  connected between themselves in order to form said section  4 . Such a fuselage  2  comprises a part  6  of its structure that is substantially cylindrical as represented on part  6 , enlarged in  FIG. 1 . 
     In order to obtain the rigid and light characteristics of fuselage  2 , panel  5  comprises a coating  7 . Such a coating  7 , also referred to as skin, is relatively thin. In addition, structural elements  8  are attached to said coating. The structural elements  9  are extended in shape and develop according to a direction substantially parallel to the generators of fuselage  2 , also referred to as longitudinal axis  100  of panel  5 . These structural elements  9  will hereinafter be referred to as stiffeners  9 . Moreover, structural elements  10  develop in a plane substantially perpendicular to the generators of fuselage  2 . These structural elements  10  will hereinafter be referred to as frames  10 . 
     In aircraft  1 , the stiffeners  9  are generally installed on an internal side of fuselage  2 . More particularly, the stiffeners  9  are attached to the skin  7 . Said skin  7  is located on an inner wall  11  of the sections  4  or panels  5 , i.e. on an inner wall of the fuselage  2 . 
     Stiffeners  9  are generally attached on the skin  7  of a first panel  13  and a second panel  14  before said panel  13  and  14  are assembled together. The stiffeners  9  attached to skin  7  of said panels  13  and  14  are therefore interrupted near to the edges  12  of said panels  13  and  14 , also referred to as ends  12  of panels  5 . The first panel  13  and the second panel  14  are assembled in such a way that their ends  12  are placed close to each other and form a line of interface  16 . However, stiffeners  9  must be substantially aligned from a first panel  13  to the second following panel  14  in order to enable the stresses of stiffener  9  to be transferred from a first panel  13  to stiffener  9  of a second panel  14 . In the state of the art, this alignment is practically impossible to achieve with acceptable levels of tolerance for battening the stiffeners without requiring long and expensive procedures. 
     In this  FIG. 1 , an area  15  of the assembly can be observed and will be given in more detail in the following figures. This area  15  corresponds to the detailed image of the elements according to the line of interface  16  taken at the junction between the first panel  13  and the second panel  14 . The stiffeners  9  are therefore substantially opposite each other on such a line of interface  16 . 
     This  FIG. 1  also shows that such lines of interface  16  of panels  5  with stiffeners  6  can exist outside of the fuselage  2 , for panels  5  of the vertical or horizontal tail units or for the wings. 
       FIGS. 2A and 2B  represent two examples of transversal cross-sections of panels comprising the (Ω) omega-shaped stiffeners. A stiffener  9  with a transversal omega-shaped cross-section comprises a first sole plate  17 , a second sole plate  18  and a body  19  of stiffener  9 . The first sole plate  17  and the second sole plate  18  are located on either side of the body  19 . The body  19  connects sole plates  17  and  18  to each other. 
     In a first mode of embodiment,  FIG. 2A , the body  19  comprises a first lateral side  20 , referred to as first core  20 , and a second lateral side  21 , referred to as second core  21 , which are connected to the first sole plate  17  and second sole plate  18  respectively. Cores  20  and  21  are connected together via a head  22 . Such a head  22  develops in a manner substantially parallel to sole plates  17  and  18 . This stiffener  9  is attached to the skin  7 . More particularly, stiffener  9  is attached to the skin  7  of panel  5  by its sole plates  17  and  18 . 
     In a second mode of embodiment,  FIG. 2B , the omega-shape can be obtained with the presence of the first sole plate  17  and second sole plate  18  connected together by a surface  23  with a rounded cross-section. Such a surface  23  performs the same role as cores  20  and  21  and as head  22  of a stiffener  9  created according to the first mode of embodiment. 
       FIG. 3  represents an overhead schematic view of the junction area of the fuselage panels according to a first mode of embodiment of the disclosed embodiments. 
     This  FIG. 3  represents the interface line  16  between the first panel  13  and the second panel  14 . Panels  13  and  14  are self-stiffened panels  5 . Panels  13  and  14  thus comprise stiffeners  9  as described above. During the assembly operations involving the first panel  13  and the second panel  14 , stiffeners  9  of the first panel  13  are located substantially opposite stiffeners  9  of the second panel  14 , with respect to the line of interface  16 . Thus, from the first panel  13  to the second panel  14 , according to a direction parallel to axis  100  of fuselage  2 , the stiffeners  9  substantially preserve the same alignment in the fuselage  2  assembly, despite the interruptions. 
     According to the disclosed embodiments, in order to ensure the continuity of the stresses between the first panel  13  and the second panel  14 , a shroud  24  is installed. Such a shroud  24  is assembled on the one hand on the first panel  13  and on the other hand on the second panel  14 . In order to lower the stresses being transmitted from panels  13  and  14  to sole plates  17  and  18 , cores  20  and  21 , in addition to head  22  of stiffeners  9  are progressively stopped before the line of interface  16 . 
     The shroud  24  covers both one part of the first panel  13  and one part of the second panel  14 . Such a shroud  24  creates a physical connection between panels  13  and  14 . Such a shroud  24  also battens sole plates  17  and  18  of stiffener  9 . Thus, the shroud  24  covers the ends  25  of sole plates  17  and  18  of stiffener  9 . Advantageously, sole plates  17  and  18  of stiffeners  9  are locally widened according to a direction substantially perpendicular to axis  100 , also referred to as the transverse direction with respect to the directions of stiffeners  9 . This widened area is located at the ends  25  of each of the sole plates  17  and  18 . These widened sole plates  17  and  18  enable the attachments of stiffener  9  to be correctly positioned, despite the possible offset between said stiffeners  9  and the two panels  13  and  14 , connected to the manufacturing tolerances. These widened sole plates  17  and  18  thus enable the shroud  24  to batten sole plates  17  and  18  while complying with the distance restrictions between the attachments and the edges of the shroud  24 . 
     According to the disclosed embodiments, on the end  12  of each panel  5 , a thickness adjusting element  26  is supported, also referred to as a doubler  26 . Such a doubler  26  performs the function of supporting the shroud  24 , the doubler  26  extending along a surface area at least equal to the surface area covered by shroud  24 , on either side of the line of interface  16  of said panels  13  and  14 . According to the longitudinal axes of the panels, parallel to axis  100 , doubler  26  extends over a distance at least equal to the distance over which shroud  24  extends. Doubler  26  acts as a regular and continuous support surface for the shroud due to the fact that the surface of the doubler of the side of shroud  24  is in continuity with the free surface of the sole plates of stiffeners  9 . 
     In a first mode of embodiment, doubler  26  is toothed, i.e. doubler  26  has the shape of a fingered plate. Toothed doubler  26  comprises a band  27  that extends along the interface  16  of panels  13  and  14 . Parallel to axis  100 , this band  27  extends over a distance  28  substantially equal to the distance separating the edge  12  of a panel  5  and an end  29  of stiffeners  9 . Moreover, this band  27  extends over the entire width of the panel  5 , according to the transverse direction. 
     Doubler  26  also extends over a surface  30  located between two neighbouring stiffeners  9 . At least one edge of doubler  26  is toothed. The toothed edge has at least one extension referred to as foot  31 , this foot  31  extending along surface  30 . The length  32  of the feet  31 , parallel to axis  100 , is adapted to the stresses being transmitted from a stiffener  9  of the first panel  13  to stiffener  9  of the second panel  14  located opposite each other, as well as to shroud  24 . Typically, length  32  added to length  28  is more than or equal to the length of shroud  24  according to a direction parallel to axis  100 . In this manner, doubler  26  acts as a continuous and regular support for shroud  24 . The width of feet  31 , according to the transverse direction, substantially covers the entire surface  30  of skin  7  between two stiffeners  9  located side by side on the same panel  5 . 
     In this mode of embodiment, feet  31  and band  27  form a single part. The thickness of doubler  26  is such that the support surface formed by sole plates  17  and  18  and by doubler  26  is regular and continuous. Typically, doubler  26  has a thickness substantially equal to the thickness of sole plates  17  and  18 . 
     Moreover, as illustrated in  FIG. 3 , frames  10  can be attached to shroud  24  at the level of the line of interface  16 . More particularly, a frame  10  is attached to shroud  24  at the line of interface  16  between the first panel  13  and the second panel  14 . Thus, according to the disclosed embodiments, no opening is created in frame  10  in order to transmit the stresses between a stiffener  9  of the first panel  13  to the second panel  14 , these stresses following a stress routing passing underneath said frame  10 , between frame  10  and coating  7 . 
     In the first mode of embodiment and in the event of a panel made out of composite material, doubler  26  is advantageously cofired with skin  7 . Such a mode of embodiment requires progressive folds. 
       FIG. 4  represents a cross-section of the first mode of embodiment of the disclosed embodiments, the cross-section plane being located on a foot of the doubler between two stiffeners  9  located side by side. 
     The presence of doubler  26  acts as a continuous and regular support at shroud  24 . Thus, doubler  26  extends over the entire surface between sole plates  17  and  18  and the two stiffeners  9  located side by side. In this mode of embodiment, the doubler is integrated into the self-stiffened panel before assembling the panels together. Doubler  26  is therefore interrupted at the level of the line of interface  16  of the two panels  13  and  14 , by virtue of the nature of the self-stiffened panels. 
     For a self-stiffened panel made out of composite material, doubler  26  is thus produced from two parts, a first part  33  of doubler  26  cofired with the first panel  13  and a second part  34  of doubler  26  cofired with the second panel  14 . 
     In order to achieve the progressive stop, a slope is created on the body  19  of stiffener  9 . This progressive stop can comprise a lip  35  substantially perpendicular to sole plates  17  and  18  for the head  22 , and a progressive slope  36  capable for example of being at sloped by approximately 45° for cores  20  and  21 . 
       FIG. 5  represents a cross-section of this first mode of embodiment of the disclosed embodiments, the cross-section plane being a plane of symmetry of a stiffener. Between the two stiffeners  9  of a first and second panel ( 13 ,  14 ), the interruption of the doubler  26  is located near to the interruption of sole plates  17  and  18 . This proximity is such that only the clearances required by the assembly tolerances are present between the doubler  26  and sole plates  17  and  18 . The progressive stop of the head  19 , having led to the distribution of the stress in sole plates  17  and  18 , the interruption of doubler  26  must be such that the stresses can easily pass from sole plates  17  and  18  to doubler  26  and from doubler  26  to shroud  24 . 
       FIG. 6  represents an overhead schematic view of the junction area of the fuselage panels according to a second mode of embodiment of the disclosed embodiments. According to this second mode of embodiment of the disclosed embodiments, doubler  26  is integrated into the skin  7 . More particularly, the doubler is thus located inside the skin  7  of a panel  5 . Such an insertion of doubler  26  to skin  7  of a panel  5  causes a variation in the level of the surface of skin  7 , shroud  24  thus being directly attached to skin  7 . 
     For a self-stiffened panel  5  made out of composite material, doubler  26  is advantageously cofired in the layers of the panel. Such a doubler causes the formation of layers and slopes  37  of said skin  7 . 
       FIG. 7  represents a cross-section of the panels between two stiffeners located side by side according to this second mode of embodiment of the disclosed embodiments. Doubler  26  is integrated into the skin  7 . As for the first mode of embodiment of the disclosed embodiments, doubler  26  is interrupted by the line of interface  16  of the first panel  13  and of the second panel  14 . 
       FIG. 8  represents an overhead schematic view of the junction area of the fuselage panels according to a third mode of embodiment of the disclosed embodiments. In contradiction to the two first modes of embodiment where doubler  26  was integrated in panels  5 , doubler  26  is an insert, i.e. manufactured independently from panels  5 . This doubler insert is installed during the assembly operations assembling panels  13  and  14  together. Moreover, doubler insert  26  is not interrupted by the line of interface  16  between the first panel  13  and the second panel  14 . 
     By virtue of its continuity, doubler  26  transfers the stresses from stiffener  9  of the first panel  13  to stiffener  9  of the second panel  14  located opposite each other. This transfer of stresses takes place in the same manner as the transfer of stresses performed by shroud  24 . 
     In a variation of the disclosed embodiments according to this mode of embodiment, doubler  26  is integrated into shroud  24 . Such a doubler  26  integrated into the shroud simplifies the operations for assembling panels  5  together. Such a part can be produced, for example, using thermosealed thermoplastic parts. 
       FIG. 9  represents a cross-section of the panels between two stiffeners located side by side according to this third mode of embodiment of the disclosed embodiments. Doubler  26  is continuous despite the line of interface  16 . In order to lower the stresses from stiffener  9  into shroud  24 , skin  7  is locally thickened. More particularly, skin  7  is thickened just before the battening area, i.e. the thickening of skin  7  is located on a part of panel  5  beginning at the edge  12  of said panel  5  and stopping further away from the line of interface than end  25  of sole plates  17  and  18 . The thickening  38  of skin  7  causes doubler  26  to be raised  39 . Moreover, this thickening  38  causes stiffener  9  to be raised  40 . Stiffeners  9  are produced to adopt the shape of the panel in such sloping areas. 
       FIG. 10  represents an overhead schematic view of the junction area of the fuselage panels according to a fourth mode of embodiment of the disclosed embodiments. In this mode of embodiment of the disclosed embodiments, doubler  26  is not an independent part. In fact, in this fourth mode of embodiment, the function performed by doubler  26  in the first three modes of embodiment is directly performed by skin  7 . Skin  7  is thus thickened in order to act as doubler  26 . 
     For a panel  5  made out of composite material, skin  7  is subjected to significant folding  42  from the edge  12  of panel  5  to an area further away from the edge of panel  5  than the beginning of the battening area. The folding  42  of skin  7  is similar to that described in the third mode of embodiment, however in this instance, the folding  42  in skin  7  of panel  5  is significant enough to act as doubler  26 . 
       FIG. 11  represents a cross-section between two stiffeners according to this fourth mode of embodiment of the disclosed embodiments. A thickness of skin  7  according to this mode of embodiment requires the role of the doubler  26  to be interrupted by the line of interface  16  as is the case for a doubler  26  integrated into the skin  7  of panel  5 . 
       FIG. 12  represents an overhead schematic view of the junction area of the fuselage panels according to a fifth mode of embodiment of the disclosed embodiments. In such a mode of embodiment of the disclosed embodiments, the role of the doubler is performed by sole plates  17  and  18  of the neighbouring stiffeners  9 . In order to achieve this, sole plates  17  and  18  are widened, according to the transverse direction with respect to an axis of the stiffeners, so that their edges are in immediate proximity to each other. Typically, a first sole plate  17  of a first stiffener  43  is locally widened according to the transverse direction. In addition, a second sole plate  18  of a second stiffener  44 , the first stiffener  33  neighbouring the second stiffener  44 , is locally widened according to the transverse direction. This widening of sole plates  17  and  18  is such that the edge of the first sole plate  17  of the first stiffener  43  is in immediate proximity to the edge of the second sole plate  18  of the second stiffener  44 , these two sole plates being separated only by the clearances required by the assembly tolerances. In the areas where stiffeners  43  and  44  are too distant from each other, in such a way that the sole plates should not be widened, this mode of embodiment can be combined with one of the other modes of embodiment previously described. The modes of embodiment of the disclosed embodiments can, as a general rule, be combined together. 
     Moreover, sole plates  17  and  18  are also extended according to a direction substantially parallel to axis  100 . In contradiction with the head of stiffener  9 , the sole plates extend up to the edge  12  of panel  5 . In addition, sole plates  17  and  18  extend according to the transverse direction in order to cover the surface  45  of skin  7  located between the edge  12  of panel  5  and the interruption of the heads  19  of stiffeners  9 . 
     The shroud  24  is thus directly supported by sole plates  17  and  18  of stiffeners  9  and by doublers  26  added as required. The transverse widening of sole plates  17  and  18  having covered surface  45  of skin  7  between heads  19  of stiffeners  9  and the edge  12  of the panel can be avoided by locally integrating doublers  26  covering said surface  45  of skin  7 . The purpose of these doublers  26  is thus to provide a regular and continuous support for shroud  24 . 
       FIG. 13  represents a cross-section between two stiffeners according to this fifth mode of embodiment of the disclosed embodiments. The advantage presented by this last mode of embodiment is that no additional part is required in the fuselage  2  in order to perform the role of the doubler  26 . However, doubler  26 , i.e. the widened and extended sole plates  17  and  18  of stiffeners  9  in this mode of embodiment, is interrupted at the edge  12  of panels  5 .