Patent Application
20150034763
This application claims priority to French Patent Application No. 13 57676 filed on Aug. 1, 2013, the entire content of which is incorporated by reference herein.
1. Technical Field
The subject matter disclosed herein relates to the field of aircraft fuselages, and in particular to that of the fastening of a fuselage frame element onto stiffeners provided on the skin of this fuselage.
The subject matter disclosed herein applies more particularly, but not exclusively, to the field of commercial aircraft.
2. Background
Aircraft fuselages are designed to resist loads induced by the pressurisation and the loads transmitted by the engines.
To this effect, fuselages usually include frames extending circumferentially, also referred to as “orbital frames”, as well as an aerodynamic skin, sometimes referred to as “self-stiffened skin”, fastened onto these circumferential frames and provided with stiffeners that usually extend longitudinally, which are fastened onto a inner face of the skin and which are generally referred to as “stringers”.
The stringers can have sections of varied types, for example as a T, I, J or Ω. In certain known configurations, the circumferential frames have a footing directly fastened onto the inner face of the self-stiffened skin, in which case the frames comprise notches/openings for the passing of the stringers. In other known configurations, the stringers are inserted between the circumferential frames and the self-stiffened skin, with the frames then passing above the stringers.
In both cases, and in particular in the latter case, attach angles commonly referred to as “clips” are arranged between the stringers, and provide the connection of the circumferential frames with the skin and/or with the longitudinal stiffeners.
However, the clips are set into place on the frame and fastened one after the other. This individual treatment of each clip makes the assembly process expensive, all the more so that the number of clips associated with each fuselage frame is very substantial. In addition, this phase of manufacturing of the fuselage generally requires the presence of many operators inside the section of fuselage concerned, due to the substantial number of fastening clips. Consequently, it can be difficult to simultaneously accomplish other assembly tasks from inside the section when the operators are installing and fastening these clips, which constitutes another factor for penalising the manufacturing time and costs.
There is therefore a need to optimise this phase of installing and of fastening clips, in order to limit the impact of it on the total cost of manufacturing the fuselage.
The subject matter disclosed herein therefore has for purpose to overcome at least partially the aforementioned disadvantages, relating to achievements of prior art.
To do this, the subject matter disclosed herein first has for object an aircraft fuselage frame element comprising a core provided with at least one through-opening intended for the passing of a fuselage stiffener, with the frame element further comprising, associated with each opening, a tab for the fastening of the frame element onto the fuselage stiffener, the tab being a single piece with the core and connected to the latter by a fold that delimits the opening.
The subject matter disclosed herein breaks with the current technique based on the use of clips fastened by riveting to the frame element, since the latter integrates tabs that fulfil an analogous function. Also, as they are already integrated into the frame element, these tabs facilitate the manufacture of the fuselage, for which the time and the costs of manufacturing are advantageously reduced.
From this stems a rationalisation of the operations, with in particular the suppression of the step of fastening clips onto the fuselage frame element. Moreover, once the frame element is installed on the other fuselage elements, these tabs can be fastened simply and quickly, which limits the number of operators that have to intervene within the fuselage. Other assembly operations can as such be conducted simultaneously in the section of fuselage concerned, with for consequence a reduction in the time and the total cost of manufacturing the fuselage. This advantage is further pronounced when such an operation of fastening of the tabs is automated, for example using robots.
The subject matter disclosed herein preferentially has one or several of the optional characteristics described hereinbelow.
Each opening can be an opening according to a direction of the height of the frame element, the fold forming the bottom of the opening and being opposite an opening portion of this opening, arranged on a footing of the frame element.
The tab is inclined by an angle of approximately 90° in relation to the core.
The fuselage frame element has a section in the general shape of a C, S, Z, I, J, L, T or Ω. Of course, this is the section in the portions of the frame element that are not provided with tabs.
The fuselage frame element comprises several openings spaced apart according to a longitudinal direction of the frame element, with a tab associated with each of the openings.
The fuselage frame element can be a thermosetting or thermoplastic material, preferably a composite material comprising a mixture of resin and fibres.
The tab is accomplished or carried out by folding a partially cut-out portion within the frame element.
The fuselage frame element extends over an angular sector of a magnitude of 30 to 120°. Other angular extents remain possible, without leaving the scope of the subject matter disclosed herein.
The fuselage frame element further comprises a stabilizer that connects the tab and the core, with this stabilizer being preferentially added onto the remainder of the frame.
The stabilizer takes the form of a section reinforcement in the shape of an L, with more preferably a triangular rib that connects the base and the branch of the L. Here, the base and the branch of the L are respectively intended to be thrust and fastened onto the core of the frame and the tab, or inversely.
The subject matter disclosed herein also has for object an assembly for aircraft fuselage comprising at least one stiffener as well as at least one fuselage frame element such as described hereinabove, with the stiffener passing through the opening of the frame element and being fastened to the latter using the tab thrust against the stiffener.
Preferably, the tab is fastened to the stiffener par rivets, bolts, welds, gluing, or polymerisation of the tab in contact with the stiffener.
Preferably, the assembly further comprises a stabilizer that connects the tab and the core of the fuselage frame element.
The subject matter disclosed herein also has for object an aircraft fuselage comprising at least one assembly such as described hereinabove, as well as a fuselage skin whereon is fastened each stiffener of the assembly.
The subject matter disclosed herein also has for object an aircraft comprising such a fuselage.
Finally, the subject matter disclosed herein also has for object a method for manufacturing an aircraft fuselage frame element such as described hereinabove, with the method comprising the following steps:
Preferably, the method further comprises a step that aims to remove, at least partially, a portion of the partially cut-out portion located on a footing of the frame element. This removal is carried out more preferably by cutting of the preferably cut-out portion. This removal can be carried out at various stages of the implementation of the method of manufacture.
Other advantages and characteristics of the subject matter disclosed herein shall appear in the detailed non-restricted description hereinbelow.
This description shall be made with regards to the annexed drawings wherein;
a to 7c diagrammatically show different steps of a method of manufacturing the fuselage frame element shown in
Referring to
The fuselage 2 comprises a skin 3 of which the inner face 3a fixedly carries fuselage frames 4, of which only one of them is shown in
Moreover, the fuselage 2 comprises a plurality of longitudinal stringers 8 which are stiffeners, taking the form of reinforcements that pass through the frames 4. All of the stringers 8 extending according to the longitudinal direction 6 are fastened to the inner face 3a of the skin, by conventional structure, such as rivets. The stringers 8 here have a transverse section in the general shape of Ω, but could have other shapes known to those skilled in the art.
Each frame 4 comprises a core 12 which is the central vertical portion shown in
The heel 14 constitutes the free end of the stiffener, opposite the end formed by the footing.
In the embodiment shown, the section of the frame 4 has the general shape of a “C” with the core 12 substantially perpendicular to the heel 14 and to the footing 16, formant respectively the opposite ends of the frame. However, angles different from 90° can be retained for certain frames 4, in particular for those located at the front end and at the rear end of the aircraft. Other general section shapes are however possible, for example as an I, Ω, etc. Hollow sections can also be considered, without leaving the scope of the subject matter disclosed herein.
As mentioned hereinabove, the fuselage 2 does not comprise conventional fastening clips, usually intended for providing the fastening of the fuselage frames 4 onto the skin 3 and/or onto the stringers 8 provided on the skin.
On the other hand, the frame element 4 integrates tabs 29 for the fastening of the stringers 8 which pass through it on the openings 31 released by these tabs.
More precisely in reference to
The tab 29 is substantially planar, inclined by approximately 90° in relation to the core 12. As shall be described hereinafter, one of the particularities of the subject matter disclosed herein resides in the fact that the tab is carried out by folding on the partially cut-out portion within the frame element.
In addition, note that the frame 4 comprises several openings 31 spaced apart according to the longitudinal direction 39 of this frame. An opening 31 is in fact provided for the passing of each stringer 8, fastened to the frame by the intermediary of the tab 29 associated with this opening. The frame 4 and each stringer 8 fastened to the latter together form an assembly 50 proper to the subject matter disclosed herein, intended to be an integral part of the fuselage 2.
In this embodiment, the tab 29 is therefore fastened onto the head of the stiffener 8, but the footing 16 could also be fastened to the legs of this same stiffener 8 against which this footing is pressing. The fastening can be carried out in an analogous manner, with rivets or similar items. As such, in the portions located between the stiffeners, the frame 4 is not connected to the skin 3, and a clearance is even preferentially provided between the inner face of the skin 3a and the footing 16. This clearance is more preferably substantially identical to the thickness of the legs of the stiffeners 8, against which the footing 16 is pressing.
The frame 4 is preferably a thermosetting or thermoplastic material, in particular a composite material comprising a mixture of reside and fibres, preferably carbon and/or glass fibres.
In this regard
Firstly in reference to
Then, as shown in
A step of polymerising the resin is then provided, during which the frame 4 is obtained by hardening under the effect of the heat. This step of polymerising is conventional, and can be implemented by any means known to those skilled in the art.
Then, a folding at 90° is carried out or made of the partially cut-out portion, in such a way that it forms the tab 29, as is shown in
For the forming of the tab 29, the portion of the partially cut-out portion located on the footing can be removed after the folding or before the latter. It can also be removed after or before the polymerisation, and, in this latter case, before or after the forming of the stack that aims to confer upon it the general shape of a C. However, if it is preferentially provided that the length of the tab corresponds to the height of the opening that opens into the core of the frame, the length of this tab could be higher, according to the extend of the removal carried out on the footing of the frame.
Once the frame 4 is obtained with the method which has just been described, it is preferentially brought inside the section of fuselage and installed on the stringers 8 already fastened to the skin 3. During this installation, the heads of the stringers 8 are therefore inserted into the openings 31, and the tabs 29 thrust against these same heads. It is then sufficient to proceed with the fastening of these tabs 29 onto the heads of the stringers, preferably in an automated manner, for example with robots. This considerably reduces the implementation time of this step of fastening of the frame onto the stringers. An analogous operation can be implemented for the possible fastening of the footing 16 onto the legs of the stringers 8.
As no operation of fastening clips onto the frame is required, the subject matter disclosed herein reduces the time and the costs of manufacturing the fuselage. The latter can indeed be subject to other assembly tasks when the frames are manufactured exteriorly to the section. In addition, thanks to a reduced encumbrance of the robots, other assembly tasks can be carried out inside the section during the fastening of the tabs on the stringers. In any case, as the number of fastenings is reduced, the same is true for the number of operators/robots present within the section in order to provide the assembly of the frames onto the stringers. Consequently, other tasks can be carried out simultaneously within this section, in order to further reduce the time and the cost of manufacturing this fuselage.
Furthermore, the subject matter disclosed herein is also advantageous in that the material removed for the formation of the openings for the passing of stringers is judiciously retained in order to form structure of fastening the frame onto these stringers. This confers a gain in material, since in the solutions of prior art, this removed material was simply discarded, without being functionalised, and in particular not intended to form all or a portion of the fastening clips.
For the fastening of the tabs 29 onto the stringer heads 8, conventional structure for fastening can be considered such as rivets, bolts, or an adhesive such as for example a thermosetting adhesive inserted between the two already hardened elements.
Alternatively, the polymerisation of the tabs 29 can be considered when they are placed in contact with the stringer heads 8. To do this, the initial polymerisation of the frame 4 is to be implemented without concerning the tabs 29, then hardened only later. In this case, the adherence to the stringers would result from the polymerisation of the tabs arranged in contact with the stringer heads 8.
Another possibility resides in the implementing of a step of copolymerisation of several frames 4 and stringers 8, by placing them in the same tool.
Of course, the various techniques for fastening mentioned hereinabove can be combined together.
The preferred technique will be the copolymerisation of several frames, tabs and stiffener elements and the skin, with the whole in the same single mould, or a technique in two steps via local welding of each tab onto its associated stiffener polymerised beforehand. When the frame is made from a thermoplastic material, identical or analogous fastening possibilities are offered. Another possibility resides however in the welding of thermoplastic tabs onto the stringer heads, preferably also carried out in a thermoplastic material. Here, the tabs are cut and then folded after the manufacture of the frame.
More generally, any method can be provided that aims for the carrying out of notches through the frame 4 after the manufacture of the latter or during its manufacture, for example before it is formed and/or before its polymerisation for the case of a thermosetting material, in such a way as to obtain at least one partially cut-out portion located between two notches. Then, a folding of the partially cut-out portion is carried out, in such a way as to release the opening and form the tab, as has been described hereinabove. The folding can also be carried out during or after the manufacture of the frame.
In reference to
Of course, various modifications can be made by those skilled in the art to the subject matter disclosed herein that has just been described, solely by way of non-restricted examples.
| Number | Date | Country | Kind |
|---|---|---|---|
| 13 57676 | Aug 2013 | FR | national |
