Classifications

• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
  • B64D15/02—De-icing or preventing icing on exterior surfaces of aircraft by ducted hot gas or liquid
  • B64D15/04—Hot gas application
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
  • B64D15/16—De-icing or preventing icing on exterior surfaces of aircraft by mechanical means
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D15/00—De-icing or preventing icing on exterior surfaces of aircraft
  • B64D15/12—De-icing or preventing icing on exterior surfaces of aircraft by electric heating
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D29/00—Power-plant nacelles, fairings, or cowlings
  • B64D29/06—Attaching of nacelles, fairings or cowlings
• • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
  • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
  • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
  • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
  • F01D25/005—Selecting particular materials
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
  • B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
  • B64D2033/0233—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes comprising de-icing means
• • B—PERFORMING OPERATIONS; TRANSPORTING
  • B64—AIRCRAFT; AVIATION; COSMONAUTICS
  • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
  • B64D33/00—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for
  • B64D33/02—Arrangements in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes

Definitions

• the present invention relates to a constituent element of an aircraft nacelle formed of a composite structure associated with a heating element and, more particularly but not exclusively, to a leading edge structure, in particular for air intake of aircraft engine nacelle.
• an aircraft engine nacelle forms the fairing of this engine and its functions are multiple: this nacelle comprises in particular in its upstream part a part commonly called "air inlet”, which has a general shape annular, and whose role is in particular to channel the outside air towards the engine.
• FIG. 1 appended hereto a section of such an air inlet in longitudinal section is schematically represented.
• This part of the nacelle comprises, in its upstream zone, a leading edge structure 1 comprising, on the one hand, a leading edge 2 strictly speaking commonly called “air intake lip”, and on the other hand part of a first internal partition 3 defining a compartment 5 in which are disposed means 6 for protection against frost, ie any means for ensuring the anti-icing and / or de-icing of the lip.
• the defrosting consists in evacuating the ice already formed, and that the anti-icing consists in preventing any formation of ice.
• the air intake lip 2 is fixed by riveting to the downstream part 7 of the air inlet, this downstream part having on its outer face a protective cover 9 and on its inner face acoustic absorption means 1 1 commonly referred to as "acoustic shroud"; this downstream part 7 of the air inlet defines a kind of box closed by a second partition 13.
• the thermal conduction of composite materials is lower than that of metallic materials, and in particular of aluminum.
• the air intake lip is subjected to a violent air flow that creates a serious risk of erosion on a composite material.
• One solution considered for overcoming the main disadvantages mentioned above proposes a leading edge formed by at least one multiaxial composite structure superimposed on the heating element intended for deicing and / or anti-icing.
• multiaxial composite structure a composite comprising fibers in the three directions, space, including reinforcing fibers passing through it in its thickness, to bind the layers of composites together.
• the present invention therefore aims in particular to provide a solution for using composite materials for the components of aircraft nacelle, particularly for leading edge structures, which does not have the disadvantages of the prior art.
• An object of the present invention is to provide a composite leading edge structure that allows effective anti-icing or de-icing, particularly in the case of electrical protection against frost means especially if these heating elements are mounted on the face internal of the air intake lip.
• Another object of the present invention is to provide a thermal conductivity leading edge structure optimized in the thickness of the structure to reduce temperature differences between the inner and outer skins of the leading edge, to increase the thermal efficiency of the lip system - means of protection against frost, and reduce the response time of temperature rise.
• leading edge structure it is also advantageous to be able to adapt the thermal conduction of the leading edge structure to its profile, that is to say its evolution along the axis lo ng itud i n a l d acel l e and rad ia l. More particularly, it is desirable to provide a leading edge structure in which the various aspects of the heat dissipation and, in particular, the direction of this heat dissipation in the leading edge structure are controlled, depending on the profile of the heat sink. leading edge and according to the important dimensions it implies.
• Another object of the present invention is to provide a composite leading edge structure with optimized thermal conduction while ensuring improved cohesion of the reinforcement within the matrix.
• This object of the invention is achieved with a constituent element of an aircraft nacelle formed of at least one composite structure and a heating element and comprising means of protection against icing, characterized in that the composite structure is matrix-reinforced by at least one material whose thermal conductivity at room temperature is greater than or equal to 800 W * m-1 * K-1 so as to ensure a transverse thermal conductivity within the nacelle element.
• Such a composite gives the component of the nacelle, which may be a leading edge structure, good thermal properties due to the presence of the doping material in the thickness of the composite structure, while ensuring a good resistance to about different impacts and erosion that it may be subjected to and while not impeding the cohesion of the fibers of the composite material within the matrix.
• the presence of the doping material in an appropriate manner within the matrix generates an increased thermal conductivity especially in the direction of the thickness of the composite structure (progressive thickness or conductivity, depending on the intended purpose), making it possible to reach an adequate temperature. for effective deicing and / or anti-icing on the outer skin of the leading edge while maintaining the resin of the composite structure below its glass transition temperature at all points and at all times.
• This increased conductivity also improves the properties of the resin of the composite structure during firing by homogenizing the temperature distribution in the material more rapidly during this operation and by greatly minimizing the thermal gradients and thus the internal stresses during the cooling of the composite. just after cooking.
• leading edge structure According to other optional features of the leading edge structure according to the invention:
• the composite structure is matrix reinforced by at least one homogenous powder so as to ensure a transverse thermal conductivity within the nacelle element;
• the composite structure is matrix reinforced with at least nanoparticles or nanotubes so as to ensure transverse thermal conductivity within the nacelle element;
• the rate a of material which is doping the matrix of the composite structure is between 1 and 50%;
• the rate a of material that boosts the matrix of the composite structure is between 50% and 90%;
• the composite structure is configured so that the material doping of the matrix of said structure changes in the thickness of said structure; the material doping of the matrix of said structure is superior in external folds of the composite structure forming the external face of the element;
• the composite structure is configured so that the particle size of the material doping the matrix of said structure changes in the thickness of said structure;
• the composite structure has a variable fiber density in the thickness of said structure
• the element further comprises an assembly material between the composite structure and the heating element, this assembly material being reinforced by at least one material whose thermal conductivity at ambient temperature is greater than or equal to at 800 W * m-1 * K-1 so as to provide transverse thermal conductivity within the nacelle element;
• the element further comprises a thermal insulator embedded in the heating element or covered by the heating element or separated from the heating element by a composite ply structure.
• the present invention also relates to a leading edge structure, in particular for an aircraft nacelle air intake, comprising a leading edge and an inner partition defining a longitudinal compartment inside this edge of the aircraft. attack housing deicing means, and / or anti-icing, the leading edge being formed of at least one composite structure and a heating element in which the leading edge is formed of an element as aforesaid .
• the present invention also relates to an air intake, which is remarkable in that it comprises a leading edge structure in accordance with the foregoing.
• Other characteristics and advantages of the present invention will emerge in the light of the description which follows, and on examining the appended figures, in which:
• FIG. 1 is a diagrammatic representation of an inlet section of a longitudinal section of the prior art (see preamble of the present description);
• FIG. 2 to 5 show cross-sectional views of different embodiments of an air intake leading edge structure according to the invention.
• a leading edge structure 1 intended in particular to be integrated with an aircraft engine nacelle air intake conventionally comprises, as previously described in the prior art, an edge Attached 2 and a partition 3 long internal interior defining a compartment intended to accommodate, in particular, means 6 for protection against icing type defrosting and / or anti-icing means.
• the means of protection against frost can be of any type. More particularly, these means may be pneumatic deicing and / or anti-icing means, electric placed in the leading edge 2 or defrosting means and / or internal anti-ice of any other type.
• the external face fe of the leading edge structure 2 is defined as the outer face, exposed to the external throttle and the internal face fi of the leading edge structure. 2 as the inner face of the structure delimiting the compartment.
• FIG. 2 there is shown a first particular embodiment of a leading edge structure 2 of the air intake lip according to the invention.
• this leading edge 2 may be structural. As explained above, this means that the leading edge 2 has a structure function, in addition to an aerodynamic function.
• the leading edge 2 has a variable thickness along its profile, and in particular, for example, a greater thickness at large curvatures and less significant at its ends.
• leading edge 2 is formed of a stack of particular layers.
• the deicing and / or anti-icing means are electric.
• This leading edge 2 comprises at least one composite structure 23 superimposed on a surface heating device.
• This heating device 30 consists of at least one electrically conductive layer 31 suitably electrically insulated by an electrical insulator 32.
• the electrical insulator 32 is formed for example by two layers 32 of elastomeric materials or composites placed on either side of the electrically conductive layer 31.
• the electrically conductive layer 31 or core 31 integrated in the air inlet lip 2 is designed as a heating element intended to provide calories to the structure of the lip 2 and contribute to eliminating the ice or maintain frost outside the outer surface fe of the lip 2 in contact with the freezing gas.
• It may comprise, in non-limiting embodiments, a resistive electrical circuit or a heating mat.
• a layer of an adhesive material 33 may optionally optionally be integrated at the interface of the composite structure 23 and the heating structure 30 as shown in FIGS. 2 and 3. Furthermore, it is also possible to optionally integrate a layer of thermally insulating material 34 with the air inlet lip structure 2.
• the thermal insulator 34 is embedded in the heating device 30 and, more particularly, placed in contact with the electrically conductive layer 31.
• FIG. 3 An alternative embodiment is illustrated in FIG. 3. This variant embodiment is identical to FIG. 2 with the following differences.
• the thermal insulation 34 is covered with the heating device 30 and, more particularly, placed in contact with a layer 32 of electrical insulation.
• the layer of an adhesive material 33 is placed at the interface of the composite structure 23 and the electrically conductive layer 31 of the heating structure 30, an electric insulation layer 32 having been removed.
• the heat insulating assembly 34 - heating device 30 is situated on the side of the internal face f 1 of the air inlet lip 2 and forms the internal skin of the air intake lip 2, the surface exposed to the freezing external gas being against the free face 23c of the composite structure 23.
• the heating structure 30 can be integrated, that is to say embedded in the thickness of the composite structure 23.
• FIG. 4 One of these alternative embodiments is illustrated in FIG. 4. This variant embodiment is identical to FIG. 3 with the following differences.
• the heating structure 30 and the thermal insulation 34 are placed in the heart of a composite structure by being covered with a composite structure 23 and 23d of one or more layers, respectively on the side of the outer face fe and the side of the internal face fi.
• FIG. 1 Another variant embodiment is illustrated in FIG. 1
• This variant embodiment is identical to FIG. 4 with the following differences. Only the heating structure 30 is placed in the heart of a composite structure by being covered with a composite structure 23 of one or more layers, respectively on the side of the outer face fe and the side of the inner face fi.
• the thermal insulator 34 forms the inner skin of the air inlet lip 2, the surface exposed to the freezing external gas being against the free face 23c of the composite structure 23.
• thermal insulation 34 and electrical insulator 32 in particular compatible materials of a composite structure.
• the heating structure 30 presented can then be disposed on the internal face fi of the air inlet lip 2, or else be integrated into the thickness of the composite structure 23, as shown more particularly in FIGS. 4 and 5.
• leading edge 2 it is also provided, or not, anti-erosion means which will be described later.
• the composite structure 23 and the anti-erosion means, if any, form the outer skin of the leading edge 2.
• this composite structure 23 is a structure formed of a reinforcing fiber reinforcement associated with a matrix that ensures the cohesion of the structure and the retransmission of the efforts towards the fibers.
• this matrix is reinforced by at least one material whose thermal conductivity at room temperature is greater than or equal to 800 Wm -1 -K -1 so as to ensure a transverse thermal conductivity within the leading edge structure.
• this reinforcement is inert from a chemical point of view with respect to the constituent fibers of the layers of the composite structure 23, 23d.
• this material may also be an electrically nonconductive material.
• this material is a diamond powder.
• a diamond material reinforcement significantly increases the transverse thermal conductivity of the composite material.
• this material may be nanoparticles or nanotubes, in particular but not exclusively of carbon material.
• It can be in powder form or in any other form of material.
• a particular embodiment of the invention is chosen for the rest of the description, namely the embodiment in which the matrix is reinforced by diamond powder.
• the composite structure 23 may be a multiaxial, monolithic, autoraidic or sandwich structure, configured to meet the thermal efficiency and structural withstand constraints of the leading edge structure 2.
• multiaxial is meant a composite comprising fibers in the three directions, space, including reinforcing fibers passing through it in its thickness, to bind the layers of composites together.
• monolithic is meant that the different plies (that is to say the layers each comprising fibers embedded in the resin) forming the composite material are bonded to each other without interposing soul between these plies .
• sandwich structure is meant a composite structure composed of two monolithic skins separated by at least one light core that can be made, in a non-limiting example, using a honeycomb structure.
• the composite structure 23 can thus be formed by a superposition of unidirectional folds (UD) and / or multidimensional (2D in particular) and oriented forming a preform.
• UD unidirectional folds
• 2D multidimensional
• the thermal conductivity of the composite structure 23 is determined as a function of the volume ratio of ⁇ -fibers and the volume level a of diamond powder which dopes the matrix.
• Acomposite ⁇ * Afibre + 1 "( ⁇ ) * (C (Adiamant + 1" (Cl) * To my trice) (1)
• Composite with A, A fib re, A and A diamond my trice being defined as the respective thermal conductivities of the composite structure 23, reinforcing fibers, diamond and matrix (usually a resin type plastic thermosetting or thermoplastic)
• the level a of diamond powder which dopates the matrix of the composite structure 23 is between 1 and 50%, preferably 3 to 40%, preferably 3 to 10%, in order to boost the composite structure and achieve an overall thermal conductivity order of magnitude equivalent to structural metal alloys, while allowing the composite structure 23 to retain the structural properties related to the matrix.
• This ga mme has the advantage of proposing a composite structure 23 whose thermal conductivity is improved while maintaining a macroscopically conventional matrix.
• the thermal conductivity of the composite structure 23 obtained is, therefore, 1 1 1 .6 Wm ⁇ -K "1
• leading edge structure 2 a thermal conductivity comparable to that of certain metals (aluminum for example).
• the level a of diamond powder that dopes the matrix of the composite structure 23 is between 50% to 90% and preferably 50 to 70%.
• this composite structure 23 exhibits an optimal mechanical behavior in compression.
• the thermal conductivity is evolutionarily defined according to the profile of the leading edge structure 2, in order to control the thermal behavior of the composite structure 23.
• the composite structure 23 is configured so that the matrix evolves and, more particularly, its doping by the material whose thermal conductivity at room temperature is greater than or equal to 800 Wm -1 -K "1 as the diamond powder evolves. in the thickness of the composite structure 23.
• the doping of the matrix is greater in the outer plies 23b of the composite structure 23, ie the folds forming the outer face fe of the leading edge structure 2.
• the matrix comprises a level of diamond powder greater than or equal to 60% in the outer plies 23b and a level of less than 50% in the other plies of the structure 23.
• the plies working in compression will be the most loaded with diamond (with potentially a granular behavior) when those working in traction will remain with a more conventional model.
• the level of reinforcing fibers may also vary in the thickness of the structure 23.
• the level of fibers may be greater in the outer plies 23b of the composite structure 23.
• certain external and / or internal plies 23a and 23a are selectively doped with diamond powder in an appropriate manner so as to have a doping distribution of the resin and the fiber ratio adapted to the mechanical stresses seen by the piece. nacelle.
• any isotope can be used.
• diamond powder particle size it is possible to choose diamond sizes of less than 10 ⁇ m, and preferably less than 5 ⁇ m, and preferably grains smaller than 3 ⁇ m.
• de-emulsifying powder of up to 0.1 ⁇ , which is low compared to fiber filament diameters generally between 4 and 10 ⁇ .
• the mixture obtained does not interfere with the cohesion of the fibers within its matrix of the composite structure 23.
• the diamond powder introduced into the atrium may consist of grains having a plurality of separate granulometries in order to maximize the filling rate of the granulate obtained.
• a doping of diamond powder comprising at least 50% of diamond grains larger than 1 ⁇ m and at least 30% of grains smaller than 1 ⁇ m, or even 30% of grains. size less than 0.5 ⁇ .
• the composite structure 23 is configured so that the particle size of the doping changes in the thickness of the structure 23.
• leading edge structure 2 it is possible to provide a second composite structure 23d, this structure being interposed between the heating structure 30 and the layer of thermally insulating material 20.
• the fibers of the reinforcement are carbon fibers, but it is also possible to use glass fibers or Keviar® (Aramid) or any other type of fiber according to the desired purpose.
• many matrices can be used such as an organic matrix or other.
• thermosetting resin such as epoxy resin, bismaleide-imide, polyimide, phenolic, or thermoplastic PPS (polyphenylene sulfide), PEEK (polyetheretherketone), PEKK (polyetherketone), etc.
• the nature of the material constituting the matrix may be different depending on the fold of the composite structure 23 considered and its position in the thickness of the structure 23 provided that the compatibility of the resins between them is respected.
• the constitutive substance of this envelope may advantageously also be doped with a material whose thermal conductivity is greater than or equal to 800 Wm "1 -K " 1 as diamond powder, to increase the conductivity.
• the adhesive material or materials used in the assembly of the lip 2 and, in particular, the adhesive material 33 used in the assembly of the composite structure 23 and the structure heating 30 can be similarly doped as well.
• the thermal conduction characteristics of the diamond of the composite structure, combined with those of the cooling core 30, are used in order to satisfy the requirements of the degreening, in particular the electric discharge, and / or the anti-ice and reduce the temperature difference between the inner skin fi and external fe of the lip 2.
• the diamond ratio in the thickness of the composite structure 23 is defined to provide transverse thermal conductivity and is adapted to dissipate the energy of the heating core 30 through the thickness of the composite structure 23.
• the thermal and mechanical properties of the leading edge structure 2 are significantly enhanced by the presence of diamond evolutionarily in the thickness of the composite structure 23.
• the temperature necessary to ensure defrosting and / or anti-icing is obtained without locally exceeding the glass transition temperature of the composite structure 23, while remaining compatible with the thicknesses necessary for the structural problem of an air inlet lip 2.
• leading edge structure 2 comprising one or more composite structures 23 as mentioned above can be provided by various manufacturing processes.
• a method of manufacturing the composite structure 23 in which is injected, by a method of injection molding type RTM (Resin Transfer Mold ing in Anglo-Saxon terms), the matrix mixture diamond powder, previously produced, in a mold containing the fibrous reinforcement.
• RTM Resin Transfer Mold ing in Anglo-Saxon terms
• the manufacturing method is a method of infusion of the RFI type (Resin Film infusion in Anglo-Saxon terms) in which the matrix-powder mixture of dyamant is blown into a fibrous preform under the pressure exerted by a flexible bladder in the direction transverse to the plane of the preform.
• RFI type Resin Film infusion in Anglo-Saxon terms
• the manufacturing process is a lay-up process for pre-impregnated fibers in which the dry matrix-diamond powder is combined with the dry fibers, and then the whole is polymerized in a subsequent step under vacuum and in an autoclave. .
• a calendered powder-matrix film having a higher or lower level of powder will be associated with one or more layers and in particular the outer surface layer 23b, the interface layer between the monolithic structure 23 and the structure of the heating element 31, and a set of preimpregnated fabric layers for producing the composite structure 23.
• the surface layer 23b is a diamond-doped thermoplastic matrix layer and the monolithic structure 23 is made by a thermosetting resin infusion or transfer process.
• the present invention is in no way limited to the embodiments described above, and any other alternative structures of composite materials doped with diamond powder could be envisaged.
• the diamond powder with already conductive metal matrices (titanium for example) whose thermal conductivity would be further increased provided that the melting and eutectic temperatures of these alloys and the melting mode (Vacuum, for example) preserves the chemical and / or crystalline integrity of the diamond powder to be dissolved.
• conductive metal matrices titanium for example
• the invention is not limited, in addition, to the leading edge structures, in particular the aircraft air intake lip, but encompasses any constituent element of an aircraft nacelle comprising at least one composite structure. associated with a heating element.

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• Engineering & Computer Science (AREA)
• Aviation & Aerospace Engineering (AREA)
• Mechanical Engineering (AREA)
• Chemical & Material Sciences (AREA)
• Materials Engineering (AREA)
• General Engineering & Computer Science (AREA)
• Laminated Bodies (AREA)
• Moulding By Coating Moulds (AREA)
• Control Of Resistance Heating (AREA)

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• 2012
  • 2012-10-09 FR FR1259599A patent/FR2996525B1/fr not_active Expired - Fee Related
• 2013
  • 2013-10-08 CA CA 2885966 patent/CA2885966A1/fr not_active Abandoned
  • 2013-10-08 EP EP13785525.0A patent/EP2906471A1/de not_active Withdrawn
  • 2013-10-08 CN CN201380052761.5A patent/CN104703879A/zh active Pending
  • 2013-10-08 RU RU2015116520A patent/RU2015116520A/ru unknown
  • 2013-10-08 BR BR112015006986A patent/BR112015006986A2/pt not_active IP Right Cessation
  • 2013-10-08 WO PCT/FR2013/052395 patent/WO2014057210A1/fr active Application Filing
• 2015
  • 2015-04-03 US US14/678,200 patent/US20150210400A1/en not_active Abandoned

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