Patent Application
20130157536
This application is entitled to and claims the benefit of German Application No. DE 10 2010 034 084.7 and U.S. Provisional Application No. 61/372,908, both filed Aug. 12, 2010, the disclosures of which, including the specification, drawings and abstract, are incorporated herein by reference in their entirety.
The invention relates to an aircraft interior lining component and a method for producing an aircraft interior lining component of this type.
Aircraft interior lining components not only serve in the optical design of the aircraft cabin but they also have the task of protecting components such as electrical lines, air- or water-conducting lines or other components of the aircraft, which are arranged behind the interior lining components, i.e. in a clearance between the interior lining components and the aircraft structure, from damage. To enable them to fulfil this protective function, the interior lining components have to have a certain mechanical strength. A further additional requirement is that, in the event that the loads acting on the interior lining components are so great that it is not possible to reliably rule out damage to the aircraft components arranged behind the interior lining components, any possible damage to these components has to be at least immediately visually noticeable through corresponding visible damage to the interior lining components. Finally, in the event of a sudden drop in pressure in the aircraft cabin—a so-called rapid decompression—the interior lining components must not endanger the passengers on board the aircraft or block emergency exits.
Interior lining components which are currently built into modern aircraft are generally constructed as sandwich components, described for example in DE 10 2006 041 787 A1, DE 10 2007 041 282 C1 or DE 10 2007 026 296 A1, which have a core and decorative layers applied to the core. It is furthermore known, for example from DE 10 2007 061 433 A1 and the previously unpublished DE 10 2009 012 015, to equip aircraft interior lining components with decompression elements which are integrated in the aircraft interior lining components and which, in the event of a sudden drop in pressure in the aircraft cabin, free a decompression opening through which air from the aircraft cabin can flow and thereby prevent damage to the aircraft interior lining components in the event of a rapid decompression.
The invention is based on the object of providing an aircraft interior lining component which, one the one hand, has sufficient mechanical strength to protect components arranged behind the aircraft interior lining component from damage and yet, on the other hand, is designed so that potential damage to the components arranged behind the aircraft interior lining component can be immediately visually noticeable through corresponding visible damage to the interior lining component and any risk to the passengers can be reliably prevented in the event of a rapid decompression. The invention is furthermore based on the object of providing a method for producing an aircraft interior lining component of this type.
This object is achieved by an aircraft interior lining component having the features of attached claims and a method for producing an aircraft interior lining component of this type, which has the features of attached claims.
An aircraft interior lining component according to the invention consists at least partially of a composite material. The composite material comprises a matrix in which first and second reinforcing fibres are embedded. The strength of an interface between a surface of the first reinforcing fibres and the matrix surrounding the surface of the first reinforcing fibres is greater than the strength of an interface between a surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres. In other words, the composite material used to produce the aircraft interior lining component according to the invention contains fibres which are embedded with varying strength in the matrix receiving the fibres. In the event of a certain mechanical load acting on the aircraft interior lining component, the different interface strengths between the first reinforcing fibres and the matrix on the one hand and between the second reinforcing fibres and the matrix on the other hand already result in a separation of the interface between the second reinforcing fibres and the matrix whilst the interface between the first reinforcing fibres and the matrix remains still intact.
The aircraft interior lining component according to the invention may consist wholly of a composite material of this type. However, it is also conceivable for only a portion or a plurality of portions of the aircraft interior lining component to be made from the composite material. For example, it is possible for only a core of the aircraft interior lining component according to the invention to consist of the composite material, which may be provided in the region of its surface with a decorative layer or a plurality of decorative layers. It is furthermore conceivable for only portions of the aircraft interior lining component behind which sensitive components of the aircraft are arranged to be made from the composite material and for other portions of the aircraft interior lining component which do not have to be as resistant to a load to be made from another material.
By selecting a different interface strength between the first reinforcing fibres and the matrix surrounding the first reinforcing fibres on the one hand and between the second reinforcing fibres and the matrix surrounding the second reinforcing fibres on the other hand, it is possible to specifically influence both the reinforcing mechanisms effected by the fibres in the composite material and the fracture behaviour of the composite material. In particular, it is possible to realise a “gradual” breakdown of the aircraft interior lining component by using a composite material to produce the aircraft interior lining component according to the invention, which contains reinforcing fibres which are embedded with varying strength in a matrix. This means that, in the event of a mechanical load acting on the aircraft interior lining component, individual regions of the composite material specifically break down so that the damage to the component is immediately visually noticeable. On the other hand, other regions of the composite material do not yet break down when this mechanical load acts on the aircraft interior lining component, which means that the component at least substantially maintains its basic structure. This ensures that the component itself is still able to at least partially fulfil its protective function for aircraft components arranged behind the component. In the event of a rapid decompression, for example, fragments of the aircraft interior lining component are furthermore prevented from endangering passengers on board the aircraft or blocking emergency exits.
Like the strength of the interface between the surface of the first reinforcing fibres and the matrix and the strength of the interface between the surface of the second reinforcing fibres and the matrix, it is possible to select and adapt the strengths of the matrix, the first reinforcing fibres and the second reinforcing fibres according to the specific application. For example, the strength of the interface between the second reinforcing fibres and the matrix may be selected so that a separation of the interface between the second reinforcing fibres and the matrix, and therefore visible damage to the aircraft interior lining component, is effected when the aircraft interior lining component is acted upon by a mechanical load which may result in possible damage to the aircraft components arranged behind the aircraft interior lining component. Furthermore, the (fracture) strength and the (fracture) toughness of the composite material, which are influenced by the strengths of the matrix, the first reinforcing fibres and the second reinforcing fibres as well as the interface strengths between the fibres and the matrix, may be adapted to the desired protective action of the aircraft interior lining component and the mechanical loads potentially acting on the aircraft interior lining component.
In the composite material used to produce the aircraft interior lining component according to the invention, the strength of the interface between the second reinforcing fibres and the matrix may be matched to the strength of the second reinforcing fibres so that the strength of the interface between the surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres is lower than the strength of the second reinforcing fibres. In other words, the strength of the interface between the second reinforcing fibres and the matrix and the strength of the second reinforcing fibres are matched to one another in such a way that, in the event of a mechanical load acting on the aircraft interior lining component, the interface between the surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres breaks down in the composite material before the second reinforcing fibres break down. This means that a separation of the interface between the second reinforcing fibres and the matrix takes place and the second reinforcing fibres slip through the matrix without the second reinforcing fibres thereby breaking down.
This ensures that a fibre frame which is formed by the second reinforcing fibres maintains its structure even when the mechanical load acting on the aircraft interior lining component has the effect of separating the interface between the second reinforcing fibres and the matrix. The second reinforcing fibres can therefore serve as a type of “safety net” which substantially maintains the basic structure of the aircraft interior lining component even when the interfaces between the second reinforcing fibres and the matrix, the matrix and/or or the first reinforcing fibres embedded in the matrix break down. Furthermore, the breakdown of the interfaces between the second reinforcing fibres and the matrix in the composite material results in the load acting on the aircraft interior lining component being redistributed to the first reinforcing fibres which, of course depending on the load level, are thus possibly specifically overloaded and can break. If the load acting on the aircraft interior lining component is indeed great enough to effect a separation of the interface between the second reinforcing fibres and the matrix, but is not sufficient to cause a breakdown of the matrix and/or the first reinforcing fibres, the separation of the interface between the second reinforcing fibres and the matrix is obvious through a deformation of the aircraft interior lining component. On the other hand, if the load acting on the aircraft interior lining component results in a breakdown of the matrix and/or the first reinforcing fibres, the damage to the aircraft interior lining component is visible as a result of visible fracture lines in the aircraft interior lining component. In each case, it is ensured that damage to the aircraft interior lining component is visually noticeable as desired.
The strength of the interface between the surface of the first reinforcing fibres and the matrix surrounding the surface of the first reinforcing fibres may be adapted to the strength of the first reinforcing fibres in such a way that a desired fracture behaviour of the composite material is achieved in the event of a mechanical load on the aircraft interior lining component. For example, the strength of the interface between the surface of the first reinforcing fibres and the matrix surrounding the surface of the first reinforcing fibres may be greater than the strength of the first reinforcing fibres. In this type of design of the composite material, the reinforcing effect of the first reinforcing fibres is used particularly effectively, i.e. the strength of the composite material is similar to the strength of the first reinforcing fibres since a mechanical load acting on the aircraft interior lining component results in a breakdown of the first reinforcing fibres before a separation of the interface between the first reinforcing fibres and the matrix takes place. High interface strengths between the first reinforcing fibres and the matrix are therefore the aim if a high maximum fracture strength of the aircraft interior lining component is desired.
However, a fibre-reinforced composite material with strong fibre/matrix interfaces may have a limited fracture toughness, of course depending on the fracture toughness of the fibres and particularly the matrix. If the aircraft interior lining component is to be notable for a high fracture toughness, it may therefore make sense to select the strength of the interface between the surface of the first reinforcing fibres and the matrix surrounding the surface of the first reinforcing fibres such that, in the event of a mechanical load on the aircraft interior lining component, a partial separation of the interface between the first reinforcing fibres and the matrix takes place before the first reinforcing fibres break down.
In the composite material used to produce the aircraft interior lining component according to the invention, the matrix may consist of a plastic material, in particular a plastic resin. Possible matrix materials include for example epoxy resins, polyester resins, vinyl ester resins and phenol resins. The plastic material forming the matrix is preferably introduced into the desired mould in liquid form, for example by casting, and then cured. The first and second reinforcing fibres may then be introduced into the plastic material which is present in the liquid state, for example during the casting procedure. This enables aircraft interior lining components of a more complex design to also be produced economically.
The first reinforcing fibres and the second reinforcing fibres may essentially be composed of different materials or may comprise different core materials, However, it is also alternatively conceivable to use first and second reinforcing fibres which are composed of the same material or comprise the same core material. Carbon fibres, glass fibres or aramid fibres may be used as first and/or second reinforcing fibres. For example, the first and/or the second reinforcing fibres may be made from a polyacrylonitrile material which is subjected to an oxidation treatment at 250 to 300° in air. Furthermore, the first and/or the second reinforcing fibres may be made from a polyacrylonitrile material which, in addition to an oxidation treatment, is subjected to a carbonising treatment at 100 to 1500° C. under nitrogen. For example, Tenax® fibres may be used as polyacrylonitrile fibres which are subjected to an oxidation treatment but are not carbonised. Pyromex® fibres may be used for example as polyacrylonitrile fibres which are subjected to both an oxidation treatment and a carbonising treatment.
To increase the interface strength between the first reinforcing fibres and the matrix, the first reinforcing fibres may be provided with a surface layer, a so-called sizing agent, of an adhesive agent which increases the strength of the interface between the surface of the first reinforcing fibres and the matrix surrounding the surface of the first reinforcing fibres. The adhesive agent may be an adhesive agent which ensures a chemical binding of the matrix to the fibre surface. However, additionally or alternatively to this, the adhesive agent may also effect an increase in the friction forces occurring between the fibre surface and the matrix, for example by roughening the fibre surface. The surface layer applied to the first reinforcing fibres is preferably adapted to the matrix material and the material of the first reinforcing fibres. For example, the first reinforcing fibres may be provided with a surface layer of a thermoplastic plastic material, such as polyurethane, a duromer plastic material or a plastic material to which a metal, for example nickel, is added.
The second reinforcing fibres are preferably not provided with a surface layer of an adhesive agent which increases the strength of the interface between the surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres. In other words, the second reinforcing fibres are preferably introduced into the matrix in the untreated state. If it is not sufficient to dispense with an adhesive agent to create a sufficiently weak interface between the second reinforcing fibres and the matrix, it is possible to also consider treating the surface of the second reinforcing fibres with a material which prevents the matrix from binding to the fibre surface and therefore specifically weakens the interface. For example, a surface layer which smooths the surface of the second reinforcing fibres and therefore reduces the friction forces occurring between the fibre surface and the matrix can be applied to the second reinforcing fibres.
The first reinforcing fibres may be embedded in the matrix in the form of a fibre fabric or fibre mesh, in one layer or in a plurality of layers, in the form of a fibre pad or in the form of individual fibres, which may be constructed as continuous fibres or as short fibres. The fibres may be orientated uniformly or randomly in all three spatial directions or they may have a preferred orientation in one spatial direction or in two spatial directions. The fibre structure formed by the first reinforcing fibres in the matrix may be selected so that the composite material exhibits the desired strength and the desired fracture behaviour.
In similar manner, the second reinforcing fibres may be embedded in the matrix in the form of a fibre fabric or fibre mesh, in one layer or in a plurality of layers, in the form of a fibre pad or in the form of individual fibres, which may be constructed as continuous fibres or as short fibres. Again, the fibres may be orientated uniformly or randomly in all three spatial directions or they may have a preferred orientation in one spatial direction or in two spatial directions. The design of the fibre structure formed by the second reinforcing fibres in the matrix is, however, preferably selected so that the second reinforcing fibres are suitable for acting as a “safety net” for the fragments of the aircraft interior lining component in the event of the matrix and/or the first reinforcing fibres breaking down as a result of a mechanical load acting on the aircraft interior lining component. The second reinforcing fibres therefore preferably form a fibre grid which extends at least two-dimensionally, preferably three-dimensionally, over the aircraft interior lining component and ensures that the basic structure of the aircraft interior lining component remains at least substantially maintained in the event of a breakdown of the first reinforcing fibres and/or the matrix.
In a method according to the invention for producing an aircraft interior lining component which consists at least partially of a composite material, the production of the composite material comprises the introduction of first reinforcing fibres into a matrix. Second reinforcing fibres are furthermore introduced into the matrix, wherein the strength of an interface between a surface of the first reinforcing fibres and the matrix surrounding the surface of the first reinforcing fibres is greater than the strength of an interface between a surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres. The aircraft interior lining component may be produced wholly or partially from the composite material. For example, after producing a core of the aircraft interior lining component which consists of the composite material, it is possible to apply a decorative layer or a plurality of decorative layers to the core consisting of the composite material.
The strength of the interface between the surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres is preferably lower than the strength of the second reinforcing fibres.
The matrix of the composite material is preferably made from a plastic material, in particular a plastic resin. Moulding of the composite material or the aircraft interior lining component preferably takes place by casting the plastic material which is in the liquid state and then curing the plastic material. The first and/or the second reinforcing fibres are preferably introduced into the matrix material whilst the matrix material is in the liquid state.
The first reinforcing fibres may comprise the same material as the second reinforcing fibres.
The first reinforcing fibres may be provided with a surface layer of an adhesive agent which increases the strength of the interface between the surface of the first reinforcing fibres and the matrix surrounding the surface of the first reinforcing fibres. Before applying the adhesive-agent surface layer, the first reinforcing fibres may be subjected to a surface pre-treatment which improves the adhesion of the adhesive-agent surface layer to the fibre surface. For example, the surface of the first reinforcing fibres may be roughened before the adhesive-agent surface layer is applied to the surface of the first reinforcing fibres.
The second reinforcing fibres are preferably not provided with a surface layer of an adhesive agent which increases the strength of the interface between the surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres. The second reinforcing fibres can essentially be introduced into the matrix material untreated. As an alternative to this, however, it is also possible to provide the second reinforcing fibres with a surface layer which reduces the strength of the interface between the surface of the second reinforcing fibres and the matrix surrounding the surface of the second reinforcing fibres. Before applying a surface layer which reduces the strength of the interface, the second reinforcing fibres may be subjected to a surface pre-treatment which improves the adhesion of the surface layer reducing the strength of the interface to the fibre surface. For example, the surface of the second reinforcing fibres may be roughened before the surface layer reducing the strength of the interface is applied to the surface of the second reinforcing fibres.
The first reinforcing fibres may be introduced into the matrix in the form of a fibre fabric or fibre mesh, in one layer on in a plurality of layers, in the form of a fibre pad or in the form of individual fibres. The first reinforcing fibres may be aligned uniformly or randomly in all three spatial directions or introduced into the matrix in such a way that they have a preferred orientation in one spatial direction or two spatial directions. In similar manner, the second reinforcing fibres may be introduced into the matrix in the form of a fibre fabric or fibre mesh, in one layer or in a plurality of layers, in the form of a fibre pad or in the form of individual fibres. The second reinforcing fibres may be aligned uniformly or randomly in all three spatial directions or introduced into the matrix in such a way that they have a preferred orientation in one spatial direction or two spatial directions. The second reinforcing fibres are preferably introduced into the matrix in such a way that a fibre structure is produced which enables the second reinforcing fibres to act as a “safety net” for the fragments of the aircraft interior lining component in the event of a breakdown of the matrix and/or the first reinforcing fibres which is caused by a mechanical load acting on the aircraft interior lining component.
If desired, prepregs, for example uni-directional (UD-)prepregs, may also be used to produce the aircraft interior lining component. The use of semi-finished prepregs is advantageous in that it is possible to manufacture semi-finished prepregs commercially with a high and constant quality. Furthermore, the prepreg already contains all the relevant components, i.e. fibres and matrix material as well as curing agents, catalysts, accelerating agents etc., so that the prepreg may be processed by curing to form a component without the addition of extra constituents. Finally, in a component which is produced using a prepreg, the resistance against the untreated fibres “slipping through” the matrix is lower.
A preferred exemplary embodiment of the invention is now explained in more detail with reference to the accompanying schematic drawings, which show
Each aircraft interior lining component 10 comprises a core 20 which consists of a fibre-reinforced composite material. A decorative layer 22 in a single-ply or multi-ply construction is applied to the core 20. Furthermore, a grid which consists for example of copper and acts as an electromagnetic shield may be applied to part of the surface or the entire surface of the aircraft interior lining component 10.
As shown most clearly in
However, the first and the second reinforcing fibres 26, 28 differ in that the second reinforcing fibres 28 are embedded in the matrix 24 as untreated fibres, i.e. they are not provided with a surface layer, whereas the first reinforcing fibres 26 have a surface layer 29 of an adhesive agent which increases the strength of an interface between the surface of the first reinforcing fibres 26 and the matrix 24 surrounding the surface of the first reinforcing fibres 26. The adhesive agent applied as a surface layer 29 to the first reinforcing fibres 26 is adapted to the matrix 24 and the first reinforcing fibres 26 and may be for example a thermoplastic plastic material or the like.
As a result of the fact that the first reinforcing fibres 26 are provided with an adhesive-agent surface layer 29 whilst the second reinforcing fibres 28 are introduced into the matrix 24 untreated, the strength of the interface between the surface of the first reinforcing fibres 26 and the matrix 24 is greater than the strength of an interface between the untreated surface of the second reinforcing fibres 28 and the matrix 24 surrounding the surface of the second reinforcing fibres 28. At the same time, the material of the second reinforcing fibres 28 and the material of the matrix 24 is selected so that the strength of the interface between the surface of the second reinforcing fibres 28 and the matrix 24 is lower than the strength of the second reinforcing fibres 28.
This means that, in the event of a mechanical load acting on the aircraft interior lining component 10, the interface between the surface of the second reinforcing fibres 28 and the matrix surrounding the surface of the second reinforcing fibres 28 breaks down before the second reinforcing fibres 28 themselves break down. At the same time, the separation of the interface between the second reinforcing fibres 28 and the matrix 24 in the event of a mechanical load acting on the aircraft interior lining component 10 has the effect of deflecting the load to the first reinforcing fibres 26. Therefore, on condition that the load acting on the aircraft interior lining component 10 exceeds the strength of the first reinforcing fibres 26, the first reinforcing fibres 26 are specifically overloaded and break.
This breakdown behaviour of the composite material gives the fracture pattern shown in
The production of the aircraft interior lining component 10 is explained below. In a first step, the first reinforcing fibres 26 are provided with the adhesive-agent surface layer which increases the interface strength. If required, the surface of the first reinforcing fibres 26 may be pre-treated, for example roughened, before the application of the adhesive-agent surface layer 29 in order to improve the adhesion of the adhesive-agent surface layer 29 to the surface of the first reinforcing fibres 26. The first reinforcing fibres 26 provided with the adhesive-agent surface layer 29 are then introduced into the matrix material which is in the liquid state.
The fibre/matrix material mix is then cast over the second reinforcing fibres 28, which are arranged in a mould and constructed as a three-dimensional fibre mesh. The surface of the second reinforcing fibres 28 is in the untreated state, i.e. an adhesive-agent surface layer is not applied to the surface of the second reinforcing fibres 28. A subsequent curing step, which may take place at an increased temperature, at a desired pressure and/or in a desired atmosphere, results in a composite material which forms the core 20 of the aircraft interior lining component 10. The core 20 is then provided with the decorative layer 20 and the copper grid layer, which is not shown in the Figures.
The aircraft interior lining component 10 shown in
As a result of this, in the event of a mechanical load acting on the aircraft interior lining component 10, the interface between the surface of the second reinforcing fibres 28 and the matrix surrounding the surface of the second reinforcing fibres 28 breaks down before the second reinforcing fibres 28 themselves break down. At the same time, the separation of the interface between the second reinforcing fibres 28 and the matrix 24 in the event of a mechanical load acting on the aircraft interior lining component 10 causes the load to be deflected to the first reinforcing fibres 26. Therefore, on condition that the load acting on the aircraft interior lining component 10 exceeds the strength of the first reinforcing fibres 26, the first reinforcing fibres 26 in the fibre mesh are specifically overloaded and break.
This gives the fracture pattern shown in
In the production of the aircraft interior lining component 10 shown in
| Number | Date | Country | Kind |
|---|---|---|---|
| 10 2010 034 084.7 | Aug 2010 | DE | national |
| Number | Date | Country | |
|---|---|---|---|
| Parent | PCT/EP2011/003943 | Aug 2011 | US |
| Child | 13763141 | US |
