Patent Application
20130230390
The present disclosure relates to a leading edge structure, in particular for an air inlet of an aircraft engine nacelle.
The statements in this section merely provide background information related to the present disclosure and may not constitute prior art.
As is known in itself, an aircraft engine nacelle forms the fairing of that engine and performs multiple functions: this nacelle in particular has, in its upstream portion, a part commonly called “air inlet,” which has a generally cylindrical shape, and the role of which is in particular to channel the outside air toward the engine.
As shown in the appended
The air inlet lip 2 is fixed by riveting to the downstream portion 7 of the air inlet, that downstream portion having a protective cowl 9 on its outer surface, and on its inner surface, acoustic absorption means 11 commonly called “acoustic shroud”; this downstream portion 7 of the air inlet defines a kind of box closed by a second partition 13.
As a general rule, all of these pieces are made from metal alloys, typically aluminum-based alloys for the air inlet lip 2 and the protective cowl 9, and titanium-based alloys for the two partitions 3 and 13. The cowl 9 can also be made from a composite material.
Such a traditional air inlet has a certain number of drawbacks: its weight is relatively high, its construction requires many assembly operations, and the presence of a large number of rivets affects its aerodynamic qualities.
To eliminate these drawbacks, one natural evolution is to replace the metal materials with composite materials.
Considerable research has been done so as to use composite materials, in particular for the leading edge structure 1.
However, this research has thus far come up against the problem of the thermal behavior of the composite materials and the consequences on the efficacy of the deicing or anti-icing systems put in place in the air inlet lip.
The thermal conduction of the composite materials is lower than that of metal materials, and in particular aluminum.
To date, it is not possible to reconcile the requirements relative to deicing and/or anti-icing of the air inlet lip 2 and those relative to the mechanical behavior of said lip 2 for a lip made from “traditional” composite materials.
In fact, it is not possible to achieve the necessary temperature on the outer skin of the lip to ensure anti-icing and/or to deice effectively, without thermally deteriorating the composite material by exceeding its vitreous transition temperature at various points.
Modifying the dimensions of the composite material, and more particularly reducing the thickness of the composite material, does not make it possible to reduce this problem.
Furthermore, such a modification also causes a decrease in the strength of the air inlet lip relative to the mechanical stresses, and the static strength and/or impact strength with respect to tools, birds or hail.
The present disclosure provides improved composite materials for aircraft leading edge structures, in particular for nacelles, that does not have the drawbacks of the prior art.
One aspect of the present disclosure is to propose a composite leading edge structure that offers effective anti-icing or deicing, in particular in the case of electric ice protection means.
It is also desirable to offer a leading edge structure that offers considerable resistance to any impacts, while continuing to perform an effective deicing and/or anti-icing function.
Another aspect of the present disclosure is to propose a leading edge structure with optimized thermal conduction in the thickness of the structure making it possible to reduce the temperature differences between the inner and outer skins of the leading edge, increase the thermal efficacy of the lip—ice protection means system, and reduce the temperature increase response time.
It is also advantageous to be able to adapt the heat conduction of the leading edge structure on its profile, i.e., its evolution along the longitudinal axis of the nacelle and radially.
It is also advantageous to propose a leading edge structure with a reduced mass.
One form of the present disclosure is achieved with a leading edge structure, in particular for an aircraft nacelle air inlet, comprising a leading edge and an inner partition defining a longitudinal compartment inside said leading edge accommodating deicing and/or anti-icing means, remarkable in that said leading edge is formed from at least one multiaxial composite structure placed on top of a heating element designed for deicing and/or anti-icing.
A multiaxial composite structure refers to a composite comprising fibers in all three spatial directions, having reinforcing fibers passing through its thickness, making it possible to connect the composite layers to each other.
Using such multiaxial composites to form a leading edge structure gives the latter good thermal properties due to the presence of the reinforcing fibers in the thickness of the composite structure, while ensuring excellent strength with respect to the various impacts it may undergo.
The presence of transverse reinforcing fibers creates a progressive thermal conductivity in the thickness of the composite structure, making it possible to be able to reach a suitable temperature for effective deicing and/or anti-icing on the outer skin of the leading edge while keeping the resin of the composite structure below its vitreous transition temperature at all points and all times.
The presence of transverse reinforcing fibers creates an improvement in the time necessary for the structure to reach the required temperatures for proper operation of the anti-icing and/or deicing system.
According to other optional features of the leading edge structure according to the present disclosure:
The present disclosure also relates to an air inlet comprising a leading edge structure according to the above.
The present disclosure also relates to a nacelle for an aircraft engine comprising an air inlet according to the above.
Further areas of applicability will become apparent from the description provided herein. It should be understood that the description and specific examples are intended for purposes of illustration only and are not intended to limit the scope of the present disclosure.
In order that the present disclosure may be well understood, there will now be described various forms thereof, given by way of example, reference being made to the accompanying drawings, in which:
The drawings described herein are for illustration purposes only and are not intended to limit the scope of the present disclosure in any way.
The following description is merely exemplary in nature and is not intended to limit the present disclosure, application, or uses. It should be understood that throughout the drawings, corresponding reference numerals indicate like or corresponding parts and features.
In all of these figures, identical or similar references designate identical or similar members or subsets of members.
A leading edge structure designed in particular to be incorporated into an air inlet of an aircraft engine nacelle traditionally comprises, as previously described in the prior art, a leading edge 2 (visible in
In another form, this leading edge 2 may be structural.
As previously explained, this means that the leading edge 2 has a structural function, in addition to an aerodynamic function.
The forces are additionally also reacted by the inner partition 3, which is sized appropriately.
In still another form, the leading edge 2 has a variable thickness along its profile, and in particular, for example, a greater thickness at significant curves and lesser at its ends.
Furthermore, the leading edge 2 is made up of a stack of particular layers.
In one form illustrated in
The core 21 integrated into the air inlet lip 2 is designed as a heating element intended to ensure the electrical conduction to allow deicing of the lip 2 and/or anti-icing protection of the latter part.
The insulation-heating mat assembly forms the outer skin of the air inlet lip 2.
It should be noted that the thicknesses of the different layers of the leading edge 2, illustrated in
The leading edge 2 also comprises a composite structure 23 placed on top of the assembly made up of the heating mat and the insulation 20.
In an alternative form of the leading edge 2, an anti-erosion layer is or is not also provided placed on top of the composite structure 23.
The composite structure 23 and the anti-erosion layer, if applicable, form the inner skin of the leading edge 2.
In one form, this multiaxial composite structure 23 is a monolithic structure.
“Monolithic” means that the different plies (i.e., the layers each comprising fibers embedded in resin) forming the composite material are alongside one another, without any core being inserted between those plies.
However, another form may provide composite structures 23 of the sandwich type.
A sandwich structure is a composite structure made up of two skins that may be multiaxial and that are separated by a core that can, in one non-limiting example, be made using a honeycomb structure.
Advantageously, this composite structure 23 is a multiaxial composite structure in the areas sensitive to ice.
It may thus be formed by a superposition of one-dimensional (UD) and/or two-dimensional (2D) plies oriented forming a preform, connected to each other by reinforcing fibers passing through them at least in their thickness.
The plies may be formed, in non-limiting examples, from an epoxy carbon or bismaleimide carbon (BMI) material.
A method for manufacturing such a multiaxial monolithic composite may consist of the dry assembly of dry fiber layers forming a preform with reinforcing fibers in the thickness to dope the thermal behavior, using a sewing or needling method. One example will be described later relative to
In another aspect using a needling method, pre-polymerized or metal composite needles are inserted.
The consolidation of the assembly thus obtained is then ensured by resin injection, using an infusion or RTM (Resin Transfer Molding) technique known in this field.
In one form, composite structures are proposed with reinforcing fibers in the thickness to dope the thermal behavior that are obtained by weaving, braiding or knitting, as illustrated in reference to
Furthermore, in another form, not illustrated, of the leading edge structure 2, it is also possible to provide a second multiaxial composite structure, that structure being inserted between the deicing mat and the layer of thermally insulating material 20.
Two forms of a multiaxial composite structure are illustrated in
In
This frame is woven by three types of fibers, i.e., fibers 231 in the warp direction, fibers 232 in the weft direction, and reinforcing fibers 233 passing through the thickness of the structure 23.
Thus, the first series 231 of fibers interlaced two by two extends toward the normal to the plane of the structure 23 and the second series 232 of fibers extends in the plane of the structure 23.
This structure 23 being multiaxial, it also comprises the reinforcing fibers 233 that crimp through all of the stacking layers of the fibers in the weft direction. The orientation of the reinforcing fibers 233 is inclined relative to the normal to the plane of the structure 23.
In one form, the incline angle is 30° and 60°.
In
This frame is woven by at least three types of fibers, including two types of fibers 234, 235 oriented in the weft and warp direction of the weaving, i.e., the plane of the structure 23 and arranged in a stack or interlacing, and fibers oriented vertically to reinforce the direction in the thickness of the structure 23 passing through the other two types of fibers 234, 235 to form a Cartesian reference.
Reinforcing fibers 236 are also added. They are arranged substantially parallel to the normal to the plane of the structure 23 to intercept the so-called Cartesian fibers.
They can thus form a series of upside down and juxtaposed U's.
Another form of a multiaxial composite structure 23 is proposed in
In this figure, a multiaxial composite structure 23 is made using a sewing method by tufting, in which the reinforcing fibers 237 have been tufted in the thickness of said structure 23.
According to the alternative form, it is possible to consider that the reinforcing fibers 233, 236 do or do not pass completely through the thickness of the composite structure 23.
They are also thermally conducting and can be made from carbon, copper or aluminum, these materials being cited as examples.
In the present disclosure, the thermal conduction characteristics of the reinforcing fibers 233, 236 of the monolithic composite structure are used, combined with those of the heating core 231, so as to meet deicing requirements, in particular electric, and/or anti-icing requirements, and reduce the temperature difference between the inner and outer skins of the lip.
The reinforcing fibers 233, 236 pass through the thickness of the composite structure 23 and form a grid of elements having an electrical conductivity that will participate in conducting heat between the inner skin and the outer skin of the lip 2.
They are thus suitable for dissipating the energy from the heating core through the thickness of the composite structure 23.
The thermal properties of the leading edge structure 2 are significantly reinforced by the physical properties of the reinforcing fibers 233, 236 in the thickness of the composite structure 23.
Progressive conductivity is thus ensured in the thickness of the composite structure 23, as well as a temperature difference between the outer skin and the inner skin of the lip 2.
With such a leading edge structure, the necessary temperature to perform deicing and/or anti-icing is obtained without locally exceeding the vitreous transition temperature of the composite structure 23, while remaining compatible with the thicknesses necessary for the structural issue of an air inlet lip 2.
A leading edge structure 2 according to the disclosure is thus capable of withstanding high thermal stresses as well as high mechanical stresses.
Furthermore, it makes it possible to reduce the mass of the air inlet lip 2.
It should also be noted that the density of the reinforcing fibers 233, 236 varies depending on the heat needs.
Of course, the present disclosure is in no way limited to the forms described above, and any other alternatives of multiaxial composite material structures may be considered.
| Number | Date | Country | Kind |
|---|---|---|---|
| 10/58931 | Oct 2010 | FR | national |
This application is a continuation of International Application No. PCT/FR2011/052475 filed on Oct. 24, 2011, which claims the benefit of FR 10/58931, filed on Oct. 29, 2010. The disclosures of the above applications are incorporated herein by reference.
| Number | Date | Country | |
|---|---|---|---|
| Parent | PCT/FR2011/052475 | Oct 2011 | US |
| Child | 13872325 | US |
