• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a fastening element (34a, 34b) for an aircraft turbojet engine nacelle, the fastening element (34a, 34b) comprising at least a first and a second subassembly (36a, 36b, 38a). , 38b), each having a first attachment zone (360a, 360b, 380a, 380b) for fixing each subassembly (36a, 36b, 38a, 38b) corresponding to a first part of a nacelle, each subassembly (36a, 36b, 38a, 38b) further comprising a second attachment zone (362a, 362b, 382a, 382b) for connecting each subassembly to a second part of the nacelle or to an element to be attached to the first part of the nacelle, the second subassembly (38a, 38b) comprising a fire resistant material.
    公开号:FR3054827A1
    申请号:FR1657635
    申请日:2016-08-08
    公开日:2018-02-09
    发明作者:Pierre Caruel
    申请人:Aircelle SA；
    IPC主号:
    专利说明:

    054 827
    57635 ® FRENCH REPUBLIC
    NATIONAL INSTITUTE OF INDUSTRIAL PROPERTY © Publication number:
    (to be used only for reproduction orders)
    ©) National registration number
    COURBEVOIE
    ©) Int Cl 8 : B 64 D 29/00 (2017.01)
    PATENT INVENTION APPLICATION
    A1
    ©) Date of filing: 08.08.16.(30) Priority: ©) Applicant (s): AIRCELLE Simplified joint stock company - FR. ©) Inventor (s): CARUEL PIERRE. @) Date of public availability of the request: 09.02.18 Bulletin 18/06. (56) List of documents cited in the preliminary search report: See the end of this brochure (© References to other related national documents: ©) Holder (s): AIRCELLE Simplified joint-stock company. ©) Extension request (s): @) Agent (s): CABINET GERMAIN & MAUREAU.
    ATTACHMENT ELEMENT FOR AN AIRCRAFT TURBOJET CARRIER AND PROPULSIVE ASSEMBLY COMPRISING SUCH AN ELEMENT.
    FR 3 054 827 - A1
    The invention relates to a fixing element (34a, 34b) for an aircraft turbojet engine nacelle, the fixing element (34a, 34b) comprising at least a first and a second subassembly (36a, 36b, 38a , 38b), each comprising a first fixing zone (360a, 360b, 380a, 380b) making it possible to fix each sub-assembly (36a, 36b, 38a, 38b) corresponding to a first part of a nacelle, each sub-assembly (36a, 36b, 38a, 38b) further comprising a second fixing zone (362a, 362b, 382a, 382b), making it possible to connect each sub-assembly to a second part of the nacelle or to an element to be fixed to the first part of the nacelle, the second sub-assembly (38a, 38b) comprising a fire-resistant material.

    i
    The present invention relates to a fastening element for an aircraft turbojet engine nacelle.
    Among aircraft, planes are known in particular which are powered by one or more turbojets each housed in a nacelle. The nacelles are generally designed to house a double-flow turbojet engine capable of generating, on the one hand, a flow of hot gases (also called primary flow) from the gas generator of the turbojet engine, and, on the other hand, a flow of cold air (said secondary flow) coming from the fan and circulating outside the turbojet engine, through an annular passage generally called secondary vein. The two streams are ejected from the turbojet engine from the rear of the nacelle.
    A turbojet engine nacelle generally has a tubular structure comprising:
    - an air intake, located upstream of the turbojet;
    - A middle section, intended to surround a fan casing of the turbojet;
    - a rear section, comprising an internal structure intended to surround the gas generator, that is to say the combustion chamber of the turbojet engine, the high pressure compressor and the turbine stages (low and high pressure), and a structure external which defines, with the internal structure, the secondary stream serving to channel the flow of cold air, the rear section being able to carry thrust reversing means;
    - an ejection nozzle, the outlet of which is located downstream of the turbojet engine.
    The nacelles are equipped with fixing elements, commonly called fittings, generally metallic, these elements being arranged for the fixing of certain elements (such as jacks, articulation hinges, locking elements, etc.) or for the fixing of certain cowling elements between them.
    Some of the fittings are arranged in areas of the nacelle which may be subject to fire in the event of fire, and must therefore resist fire. To this end, the fittings are made of titanium or stainless steel, which represents a high mass and cost.
    The object of the present invention is to remedy the drawbacks of the state of the art by making it possible to produce fire-resistant fittings which are light and less expensive to produce.
    To this end, it relates to a fixing element, a fixing element for an aircraft turbojet engine nacelle, the fixing element comprising at least a first and a second subassembly, each comprising a first fixing zone making it possible to fix each subassembly. corresponding to a first part of a nacelle, each subassembly further comprising a second fixing zone, making it possible to connect each subassembly to a second part of the nacelle or to an element to be fixed to the first part of the nacelle, the second sub-assembly comprising a fire-resistant material.
    Thus, by providing a fastening element comprising two sub-assemblies, one of these sub-assemblies being fire-resistant, it is possible to produce the other sub-assembly in a lighter and / or less expensive material, for example a composite material with organic matrix. The mass and the cost price of the assembly are thus greatly reduced, while ensuring the necessary resistance, in particular fire resistance.
    In one embodiment, the first sub-assembly is designed to resist so-called limit and ultimate loads.
    In one embodiment, the second sub-assembly is designed to resist so-called fatigue loads.
    In one embodiment, the second sub-assembly comprises a metallic material.
    In one embodiment, the second sub-assembly is made of stainless steel or titanium.
    In one embodiment, the first sub-assembly is made of aluminum alloy or of an organic matrix composite material.
    In one embodiment, the second sub-assembly is at least partially nested in a cavity of the first sub-assembly.
    In one embodiment, the first fixing zone of each subassembly is intended to be fixed to a part of the nacelle by means of mechanical fixing elements, the first fixing zone of the second subassembly comprising locations for fixing elements independent of the first subset.
    In one embodiment, the first and second subsets have locations for common fasteners.
    In one embodiment, the second sub-assembly is produced from the folding of at least one sheet.
    The invention also relates to a fixing assembly comprising two fixing elements in accordance with the above, the two fixing elements being connected in an articulated manner by means of their second fixing zone.
    The invention further relates to an aircraft propulsion assembly comprising a turbofan engine and a nacelle, said propulsion assembly comprising at least one attachment element as defined above, and / or at least one attachment assembly as defined above.
    The present invention will be better understood on reading the following detailed description, made with reference to the accompanying drawings, among which:
    - Figure 1 shows an aircraft propulsion unit, comprising a nacelle and a turbofan engine;
    - Figure 2 shows a cowling of the nacelle of Figure 1, comprising two half-covers connected by fastening elements according to the invention;
    - Figure 2a is a detail view of Figure 2 showing one of the fixing elements;
    - Figures 3a to 3c show an embodiment of a fastening element according to the invention.
    FIG. 1 shows a nacelle 1 equipping a turbofan engine with mixed mixed flow, in the example a low thrust turbojet engine intended in particular for business aviation. The nacelle 1 comprises in particular an air inlet section 2, located upstream of a fan of the turbojet engine, a middle section 3 surrounding the fan of the turbojet engine, and a rear section 4, located downstream of the middle section. The rear section 4 includes in particular a thrust reversing device.
    FIG. 2 shows a view of the cover of the middle section 3. The middle section 3 comprises two half-covers 30, 32, in the example an upper half-cover 30 and a lower half-cover 32. The two half-covers 30, 32, are connected to one another at one of their ends (at a position called "3 o'clock", by means of a plurality of fixing assemblies according to the invention In the example of FIG. 2, the fixing assemblies 34 constitute hinges allowing the rotation of the half-covers 30, 32 relative to each other.
    FIG. 2a is a detailed view of FIG. 2, showing more particularly one of the fixing assemblies 34. Each fixing assembly 34 comprises two half-parts 34a, 34b, each constituting a fixing element 34a, 34b conforming to the invention. As can be seen in FIG. 2a, the fastening assembly 34 comprises a fastening element 34a, of the female type, secured to the half-cover 30, and a second half-part 34b, of the male type, secured to the half-cover 32. Each of the two fastening elements 34a, 34b is attached to the corresponding half-cover 30, 32, in the example by means of a plurality of mechanical fasteners (not shown), preferably not removable, such as rivets. In addition, the two fastening elements 34a, 34b are connected to each other by means of a bolt 322, which thus makes it possible, in collaboration with the other fastening assemblies 34, to secure the two half-cowls 30, 32 by an articulated connection.
    The fixing assembly 34 is better visible in Figures 3a to 3c, which are respectively perspective, exploded and partial views of this element.
    As can be seen in FIGS. 3a and 3b, each fastening element 34a, 34b, comprises two subassemblies, namely a first subassembly 36a, 36b and a second subassembly 38a, 38b. Each of the subsets 36a, 38a, 36b, 38b is configured to be fixed, on the one hand, to the corresponding half-cap, and, on the other hand, either to an engine or nacelle component (such as a jack), or , as in the example in the figures, to another fixing element to form a fixing assembly according to the invention.
    For this purpose, each sub-assembly has two attachment zones: a first attachment zone 360a, 380a, 360b, 380b, allowing the attachment of the sub-assembly to a support such as an internal surface of the nacelle cover, and a second fixing zone 362a, 382a, 362b, 382b, allowing the subassembly to be fixed to a nacelle (or engine) component, such as a jack, or to another fixing element. In the example of FIGS. 3a to 3c, the first fixing zone comprises two flat zones provided with through holes 324, 326 allowing the passage of fixing elements such as rivets. The second fixing zone has two through holes, allowing the passage of a fixing element (such as bolt 322).
    According to the invention, the second sub-assembly 38a, 38b is made of a fire-resistant material, that is to say having fire resistance at least equal to that required by the regulations in force for certification. Thus, the second sub-assembly 38a, 38b must resist for at least 15 minutes by being subjected to a flame of 2000 K, while being subjected to forces equivalent to the forces undergone in normal operation. In the example, it is a metallic material, such as steel (especially stainless steel) or titanium. The second sub-assembly is thus able to withstand a fire-type incident occurring during operation of the turbojet engine of the propulsion assembly in which the corresponding fixing element is mounted. Furthermore, the second sub-assembly is designed to withstand at least the so-called fatigue loads (as defined by the regulations in force), which are the only loads likely to be exerted on the propulsion unit after an incident of the type fire, since the fire event is not combined with another exceptional event. Thus, the second sub-assembly is not designed to withstand standard loads of higher intensity, namely the so-called limit loads (which are loads occurring approximately ten times in the life of the aircraft), and the ultimate loads (loads occurring for example during a loss of dawn of the turbojet fan). By thus dimensioning the second subset so that it withstands only the weakest typical loads, namely the fatigue loads, which are generally two to ten times lower than the ultimate loads, its mass is greatly reduced and also the manufacturing cost. In order to further reduce the manufacturing cost, the second sub-assembly 38a, 38b can advantageously be produced by folding at least one sheet.
    The first sub-assembly 36a, 36b is advantageously made of a lighter and / or less expensive material than the second sub-assembly, for example aluminum or a composite material with an organic matrix (preferably with a reinforcement based on carbon fiber). The first sub-assembly is designed to withstand limit loads and ultimate loads, but is not necessarily fire resistant. Thus, after a fire incident, if the resistance of the first sub-assembly is greatly altered, even if it is destroyed, then the necessary resistance will be ensured by the second sub-assembly.
    FIG. 3c is a partial view showing the cooperation between the second subsets 38a, 38b. These two sub-assemblies are connected by the bolt 322, each being moreover connected to one of the two half-covers 30, 32. Thus, if the first sub-assemblies 36a, 36b are destroyed or badly damaged, it can be seen that the connection between the two half-covers is entirely ensured by the second sub-assemblies 38a, 38b, these being fire-resistant.
    The second subsets 38a, 38b advantageously comprise two half-parts 384a, 386a, 384b, 386b, symmetrical, produced by bending and drilling a metal sheet. Each of these half-portions carrying a fixing zone 380a, 380b, these zones being provided with holes 324 as mentioned above. Each half-part has a wall 388a, 390a, 388b, 390b forming a projection extending (in particular substantially perpendicularly) from the corresponding fixing zone. The walls 388a, 390a, 388b, 390b each carry a fixing orifice 382a, 382b. The projecting walls 388a, 390a, 388b, 390b are designed to be at least partially nested in a cavity or a corresponding space of the first subsets 36a, 36b, so that the orifices 362a, 362b and 382a, 382a coincide ( that is to say that their axes are combined). We can also plan to make the first sub-assemblies in two half-parts, as shown in Figure 3b for the sub-assembly 36b. It may also be envisaged that the first subsets are at least partially nested in a corresponding cavity of the second subsets.
    As can be seen in FIGS. 3a and 3b, part of the holes 324 can be made to coincide with certain holes 326, the holes of coincident axes then being crossed by a common fixing element. This arrangement allows in particular the first sub-assembly to protect from the flame, at least partially and during part of the exposure time, the fixing zone in the corresponding half-cover.
    So that the different sub-assemblies have the closest possible expansion, it is possible, for example, to associate a titanium sub-assembly with a composite material sub-assembly, and a steel sub-assembly with an aluminum sub-assembly.
    Although the invention has been described with a particular embodiment, it is obvious that it is in no way limited thereto and that it includes all the technical equivalents of the means described as well as their combinations if these fall within the scope of the invention.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    1. Fastening element (34a, 34b) for an aircraft turbojet engine nacelle, the fastening element (34a, 34b) comprising at least a first and a second subassembly (36a, 36b, 38a, 38b), each comprising a first fixing zone (360a, 360b, 380a, 380b) making it possible to fix each sub-assembly (36a, 36b, 38a, 38b) corresponding to a first part of a nacelle, each sub-assembly (36a, 36b, 38a, 38b) further comprising a second fixing zone (362a, 362b, 382a, 382b), making it possible to connect each sub-assembly to a second part of the nacelle or to an element to be fixed to the first part of the nacelle , the second subset (38a, 38b) comprising a fire-resistant material.
    [2" id="c-fr-0002]
    2. Fastening element (34a, 34b) according to one of the preceding claims, in which the second sub-assembly (38a, 38b) comprises a metallic material.
    [3" id="c-fr-0003]
    3. Fastening element (34a, 34b) according to the preceding claim, wherein the second sub-assembly (38a, 38b) is made of stainless steel or titanium.
    [4" id="c-fr-0004]
    4. Fastening element (34a, 34b) according to one of the preceding claims, in which the first sub-assembly (36a, 36b) is made of aluminum alloy or of composite material with organic matrix.
    [5" id="c-fr-0005]
    5. Fastening element (34a, 34b) according to one of the preceding claims, in which the second sub-assembly (38a, 38b) is at least partially nested in a cavity of the first sub-assembly (36a, 36b).
    [6" id="c-fr-0006]
    6. Fastening element (34a, 34b) according to one of the preceding claims, in which the first fixing zone (360a, 360b, 380a, 380b) of each sub-assembly (36a, 36b, 38a, 38b) is intended to be fixed to a part of the nacelle by means of mechanical fixing elements, the first fixing zone (380a, 380b) of the second sub-assembly (38a, 38b) comprising locations for fixing elements independent of the first sub -set (36a, 36b).
    [7" id="c-fr-0007]
    7. Fastening element (34a, 34b) according to one of the preceding claims, in which the first and second subassemblies (36a, 36b, 38a, 38b) have locations for common fastening elements.
    [8" id="c-fr-0008]
    8. Fastening element (34a, 34b) according to one of claims 2 to 7, wherein the second sub-assembly (38a, 38b) is made from the folding of at least one sheet.
    [9" id="c-fr-0009]
    10 9. Fastening assembly (34) comprising two fastening elements (34a, 34b) according to one of the preceding claims, the two fastening elements being connected in an articulated manner by means of their second fastening zone.
    [10" id="c-fr-0010]
    15 10. Aircraft propulsion assembly comprising a turbofan engine and a nacelle, said propulsion assembly comprising a fixing element (34a, 34b) according to one of claims 1 to 8, or a fixing assembly (34) according to claim 9.
    1/4
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    同族专利:
    公开号 | 公开日
    FR3054827B1|2019-08-23|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    DE678518C|1936-11-05|1939-07-17|Henschel Flugzeug Werke A G|Clamping device for cladding of aircraft, especially for engine hoods|
    GB2288578A|1994-04-18|1995-10-25|Short Brothers Plc|An aircraft propulsive power unit|
    US20090173823A1|2008-01-07|2009-07-09|Rohr, Inc.|Method and component for determining load on a latch assembly|
    US20110062279A1|2009-09-11|2011-03-17|Spirit Aerosystems, Inc.|Hybrid beam for a thrust reverser unit|
    EP2436601A1|2010-10-01|2012-04-04|Airbus Operations|Thrust link device with connecting rods for an aircraft engine mounting, including three aligned ball-and-socket joints|
    EP2546148A2|2011-07-14|2013-01-16|Hamilton Sundstrand Corporation|Fire resistant structural mount yoke and system|
    EP2706011A1|2012-09-06|2014-03-12|Airbus Operations |Lateral propulsion assembly for an aircraft including a turboshaft engine support arch|
    FR3013682A1|2013-11-27|2015-05-29|Rohr Inc|
    FR3015433A1|2013-12-23|2015-06-26|Airbus Operations Sas|AIRCRAFT ASSEMBLY COMPRISING AN INTEGRATED PLATFORM LOADING MACHINE AND REAR PARTLY FUSELING AGENCY|EP3549856A1|2018-04-03|2019-10-09|Rohr, Inc.|Fan cowl latch concept for fuselage mounted power plant|
    EP3647203A1|2018-11-05|2020-05-06|Rohr, Inc.|Nacelle cowl hinge|
    EP3800126A1|2019-10-04|2021-04-07|Rohr, Inc.|Cowl door latch assembly|
    法律状态:
    2017-07-06| PLFP| Fee payment|Year of fee payment: 2 |
    2018-02-09| PLSC| Search report ready|Effective date: 20180209 |
    2018-08-30| PLFP| Fee payment|Year of fee payment: 3 |
    2019-07-22| PLFP| Fee payment|Year of fee payment: 4 |
    2020-07-21| PLFP| Fee payment|Year of fee payment: 5 |
    2021-07-22| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1657635A|FR3054827B1|2016-08-08|2016-08-08|FIXING ELEMENT FOR AIRCRAFT TURBO BURNER BOOM AND PROPULSIVE ASSEMBLY COMPRISING SUCH A COMPONENT|
    FR1657635|2016-08-08|FR1657635A| FR3054827B1|2016-08-08|2016-08-08|FIXING ELEMENT FOR AIRCRAFT TURBO BURNER BOOM AND PROPULSIVE ASSEMBLY COMPRISING SUCH A COMPONENT|
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