• 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 thermal protection system (58) of the rigid structure (10) of a pylon (4) in a propulsion unit (1) of known type of aircraft induces an additional penalty weight and occupies a specific volume that 'must be taken into account when designing the propulsion system. To remedy these problems, the invention proposes to house the duct (60) of the thermal protection system within the box (24) of the rigid structure, so as to benefit from the volume allocated to the rigid structure (10). The elements (26, 62, 30) forming the ducts (60) can thus be an integral part of the rigid structure and can therefore play a structural role.
    公开号:FR3020798A1
    申请号:FR1454183
    申请日:2014-05-09
    公开日:2015-11-13
    发明作者:Wolfgang Brochard；Misael Hernandez-Perez；Yves Belloc
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD The present invention relates to the field of engine attachment poles intended to be interposed between an aircraft wing and an airfoil. The invention relates to the field of engine attachment poles intended to be interposed between an aircraft wing and an airfoil. a motor, and more particularly to a propulsion unit comprising such a mounting pylon.
    [0002] The invention can be used on any type of aircraft equipped for example with turbojets or turboprops. This type of attachment pole, also called "EMS" (of the English "Engine Mounting Structure"), allows for example to suspend an aircraft engine below the wing of the aircraft, or to mount such a motor above the wing. STATE OF THE PRIOR ART In general, such an attachment pylon is provided to form the connecting interface between an aircraft engine and the wing of the aircraft. It makes it possible to transmit to the structure of the aircraft the forces generated by the engine, and also authorizes the routing of fuel, electrical, hydraulic and air systems between the engine and the airframe of the aircraft. To ensure the transmission of forces, the attachment mast has a rigid structure, also called "primary structure", which is generally of the "box" type, that is to say formed by the assembly of longitudinal members upper and lower closure and two panels or closing side walls. These elements are generally connected to each other via transverse stiffening ribs, when the dimensions of the box justify it.
    [0003] On the other hand, the attachment mast is provided with a mounting system interposed between the engine and the rigid structure of the suspension pylon. This mounting system generally comprises at least two engine attachments, generally a front attachment and a rear attachment.
    [0004] In addition, the mounting system generally comprises a device for taking up the thrust forces generated by the engine. In the prior art, this device takes for example the form of two lateral rods connected on the one hand to a rear part of the hub of the intermediate casing arranged in the continuity of the turbojet fan casing, or to a front part of the surrounding casing. the heart of the turbojet, and secondly to the rear engine attachment fixed on the housing of the latter. Similarly, the attachment mast also comprises a second mounting system interposed between the rigid structure of the mast and the wing of the aircraft, the second system is usually composed of several fasteners. Finally, the attachment mast is provided with a secondary structure ensuring the segregation and maintenance of the systems while supporting aerodynamic fairings, the rear aerodynamic fairing usually protruding beyond the trailing edge of the wing, to the rear. Furthermore, the propulsion unit is generally equipped with a thermal protection system to protect the box of the rigid structure of the attachment pylon vis-à-vis the heat released by the engine. Such a thermal protection system comprises a conduit arranged under the box to allow the circulation of a relatively cool air flow between the box and the engine, so as to maintain a moderate temperature within the primary structure, especially in which concerns the systems housed within this structure.
    [0005] However, the arrangement of this conduit under the box results in the need to allocate a specific volume to the duct during the design and sizing of the propulsion unit. Now it is desirable for aircraft designers to be able to overcome this constraint. In addition, such an arrangement of the conduit leads to an excess of mass which is undesirable.
    [0006] The international application WO 2007/036521 of the Applicant describes an example of a known type of propulsion assembly illustrating the problem explained above. DISCLOSURE OF THE INVENTION The invention aims in particular to provide a simple, economical and effective solution to these problems, to avoid at least partly the aforementioned drawbacks. The invention proposes for this purpose a propulsion unit for an aircraft, comprising a motor and an attachment pylon for attaching the engine to the wing of an aircraft, the pylon comprising: a rigid structure comprising a box having a first closure beam, a second closure beam opposite the first closure beam, and two closure side walls each having a first end attached to the first closure beam and a second end attached to the second closure beam; a mounting system ensuring attachment of the engine to the box, and a thermal protection system comprising a conduit forming a thermal barrier to protect the box vis-à-vis the heat of the engine. According to the invention, said conduit is housed inside said box.
    [0007] Thus, the interior volume of the box is used to house the duct of the thermal protection system, so that it no longer requires specific volume to consider in the design of the propulsion system. Due to the integration of the elements forming the conduit to the rigid structure, these elements can be used to participate in the transfer of the thrust and support forces of the engine to the wing of the aircraft. These elements can thus have a structural function in addition to their function of delimiting an air circulation channel. In a preferred embodiment of the invention, said box comprises an inner spar having lateral ends connected to said closing side walls, said inner spar extending between said first closure spar and said second closure spar, and said duct is delimited by the first closure beam, by the inner spar, and by the closing side walls.
    [0008] Preferably, said mounting system comprises a rear engine attachment, and said rigid structure of the attachment pylon comprises a structural block comprising: a mounting plate applied to said first closure spar and fixed thereto by means of members fastener passing through the first closure beam and said inner spar, and - an attachment interface of said rear engine attachment. Preferably, said thermal protection system comprises spacers connecting said first closure beam to said inner spar passing through said conduit, each spacer incorporating at least one orifice for the passage of one of said fasteners of said structural block. Preferably, a first region of said duct, facing said structural block, has an enlarged cross section in the direction from said first closure spar to said inner spar. Said enlarged cross-section of said first region of the duct is advantageously dimensioned so that the effective cross-section of air passage is constant along said duct. Preferably, a second region of said duct, traversed by an air supply duct intended to be connected to the cell of an aircraft, has a widened cross section in the direction from said first closure spar to said inner spar. Said enlarged cross-section of said second region of the duct is advantageously sized so that the effective cross-section of air passage is constant along said duct.
    [0009] Preferably, said mounting system comprises a front engine attachment having a fixing plate applied to said first closure spar and fixed thereto. Preferably, said thermal protection system comprises a longitudinal partition wall which extends between said first closure spar and said inner spar and which is connected to said closing side walls, so as to share a central region of the duct in two parts. . Preferably, said box comprises a front closure rib connecting a front end of said first closure beam to a front end of said second closure beam. Preferably, said box comprises a rear closing rib connecting a rear end of said first closure beam to a rear end of said second closure beam. Preferably, said box comprises internal stiffening ribs having an end connected to said inner spar and an opposite end connected to said second closure spar. Preferably, said thermal protection system comprises at least one air inlet connected to said duct and opening through one of said closing side walls or through said first closure beam. Preferably, said thermal protection system comprises at least one air outlet connected to said duct and opening through one of said closing side walls or through said first closure beam. Preferably, said thermal protection system comprises longitudinal stiffening ribs against buckling, said longitudinal ribs forming heat exchange fins. The invention also relates to an aircraft comprising at least one propulsion unit of the type described above. BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood, and other details, advantages and characteristics thereof will appear on reading the following description given by way of nonlimiting example and with reference to the appended drawings in which: - Figure 1 is a partial schematic view in axial section of a propulsion assembly according to a preferred embodiment of the invention; - Figure 2 is a partial schematic view in axial section of a box belonging to a rigid structure of a pylon of the propulsion assembly of Figure 1; FIG. 3 is an enlarged view of detail IV of FIG. 2 illustrating a front end portion of the rigid structure of the attachment pylon; - Figure 4 is a partial schematic perspective view of the front end portion of the rigid structure of the attachment pylon; - Figure 5 is a partial schematic perspective view of an alternative embodiment of the front end portion of the rigid structure of the attachment pylon; FIG. 6 is an enlarged view of detail VII of FIG. 2 illustrating a rear end portion of the rigid structure of the attachment pylon; - Figure 7 is a partial schematic exploded perspective view of the rigid structure of the attachment pylon of the propulsion unit of Figure 1; FIG. 8 is a partial diagrammatic cross-sectional view of the rigid structure of the attachment pylon of the propulsion unit of FIG. 1, illustrating a first region of the box of FIG. 2, facing the structural block; - Figure 9 is a partial schematic cross sectional view of an alternative embodiment of the rigid structure of the attachment pylon of the propulsion unit of Figure 1; - Figure 10 is a partial schematic cross-sectional view of the box of Figure 2; FIG. 11 is a partial schematic cross-sectional view of an alternative embodiment of the box of FIG. 2.
    [0010] In all of these figures, identical references may designate identical or similar elements. DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS FIG. 1 illustrates an aircraft, more particularly a propulsion unit 1 fixed under a wing 3 of this aircraft. Overall, the propulsion unit 1 comprises a motor 2 such as a turbojet and an attachment pylon 4, the latter comprising in particular a rigid structure 10 and a mounting system 11 composed of a plurality of engine fasteners 6, 8 and a device for taking up thrust forces 9. The mounting system 11 is interposed between the engine 2 and the rigid structure 10. As an indication, the propulsion unit 1 is intended to be surrounded by a nacelle ( not shown in this figure), and the attachment pylon 4 comprises another series of fasteners (not shown) to ensure the suspension of the propulsion unit 1 under the wing 3 of the aircraft. In the following description, X is the longitudinal direction of the attachment pylon 4 which is also comparable to the longitudinal direction of the turbojet engine 2, this X direction being parallel to a longitudinal axis 5 of the turbojet engine 2. On the other hand, Y is the direction transversely oriented with respect to the mast 4 and also comparable to the transverse direction of the turbojet 2, and Z is the vertical direction or the height, these three directions X, Y and Z being orthogonal to each other. On the other hand, the terms "front" and "rear" are to be considered in relation to a direction of advancement of the aircraft encountered following the thrust exerted by the turbojet engine 2, this direction being represented schematically by the arrow 7. The turbojet engine 2 has, at the front, a large fan casing 12 delimiting an annular fan duct 14 and has a smaller central casing 16 enclosing the turbojet engine core towards the rear. Finally, the central casing 16 extends towards the rear by an ejection casing 17 of larger size than that of the central casing 16. The housings 12, 16 and 17 are of course integral with each other. As can be seen in Figure 1, the plurality of engine fasteners is constituted by a front engine attachment 6 and a rear engine attachment 8 for example formed of two rear half-fasteners, in a manner well known. The thrust force recovery device 9 takes for example the form of two lateral rods (only one being visible due to the side view) connected on the one hand to a rear part of the hub of the intermediate casing arranged in the continuity of the fan casing 12, and secondly to a spreader 20, itself mounted on the rigid structure 10. The front engine attachment 6, secured to the fitting 15 of the rigid structure 10 and to the fan casing 12, is conventionally designed so as to take only forces generated by the turbojet engine 2 along directions Y and Z, and therefore not those exerted in the direction X. As an indication, this front attachment 6 preferably enters a portion of circumferential end of the fan casing 12. The rear engine attachment 8 is generally interposed between the ejection housing 17 and the rigid structure 10 of the mast. As indicated above, it is preferably designed so as to be able to take up forces generated by the turbojet engine 2 along directions Y and Z, but not those acting in direction X. In this way, with the mounting system 11 of isostatic nature, the recovery of the forces exerted in the direction X is carried out using the device 9, and the recovery of forces exerted along directions Y and Z is carried out jointly with the help of the front attachment 6 and rear attachment 8.
    [0011] On the other hand, the recovery of the moment exerted in the direction X is carried out vertically with the aid of the fastener 8, the recovery of the moment being exerted in the direction Y is carried out vertically with the aid of the rear fastener 8 together with the fastener 6, and the recovery of the moment being in the direction Z is carried transversely with the aid of the fastener 8, together with the fastener 6.
    [0012] Still with reference to Figure 1, the rigid structure 10 has a box 24 extending from one end to the other of the rigid structure 10 in the X direction, and thus forms a torsion box, sometimes referred to as "main box ", Of the structure. This box 24 is formed by a first closure beam 26, also called "lower beam" when the propulsion unit is intended to be mounted under the wing as in the example shown, a second longitudinal beam 28, also called "spar in this case, as well as by two lateral closing walls 30 (only one being visible in FIG. 1) each extending in the direction X, substantially in a plane XZ. In the case of a propulsion unit intended to be mounted above the wing, the first closure spar 26 forms an "upper spar" while the second spar 28 forms a "lower spar". Inside the box 24, transverse stiffening ribs 32, arranged substantially in YZ planes and spaced longitudinally, reinforce the rigidity of the box 24. As an indication, the elements 26, 28 and 30 can each be made of in one piece, or alternatively each formed of an assembly of adjacent panels, which may optionally be slightly inclined relative to each other. As shown in Figure 1, the rigid structure 10 further comprises a structural block 34 fixedly mounted on the outer surface of the first closure beam 26. The structural block 34 has an attachment interface 36 of the rear engine attachment 8, this The interface 36 is thus situated below the plane in which the beam 28 is located. This fixing interface 36 comprises, for example, yokes 38 on which are articulated shackles also intended to be articulated on integral fittings of the engine, in a manner that known in itself. The shackle assembly then forms the rear engine attachment 8. In addition, the structural block 34 comprises a mounting bracket 20 mounted preferably forward with respect to the attachment interface 36. Referring to Figure 2 , the attachment mast 4 comprises a thermal protection system 58 of the box 24 mainly comprising a conduit 60 housed inside the box 24. This conduit 60 is intended to be traversed by a relatively cool air flow so as to form a thermal barrier against the warmer gases from the engine. The duct 60 is delimited by the first closure spar 26 and by an internal spar 62. This inner spar 62 has lateral ends connected to the closing side walls 30 and extends parallel to the first spar 26, between the first spar closure member 26 and the second closure beam 28. In addition, the duct 60 is delimited laterally by each of the closing side walls 30.
    [0013] Furthermore, the thermal protection system 58 comprises a longitudinal partition wall 64 which extends between the first closure spar 26 and the inner spar 62. This longitudinal partition wall 64 is connected, by its lateral ends, to the side walls. closure 30, so as to share a median region of the conduit 60 in two parts 66a, 66b. The longitudinal partition wall 64 makes it possible in particular to split the relatively fresh air flow into two air gaps, so as to increase the efficiency of the thermal barrier. The division of the two-part air flow also makes it possible to provide redundancy so that, in the event of an anomaly, such as a partial or total obstruction of one of the parts 66a, 66b of the duct 60, The other part continues to provide the thermal protection function of the system 58. Such redundancy is commonly referred to as a "safe function". As also shown in FIG. 2, the duct 60 has a first region 70 situated opposite the structural block 34 and having a widened cross section in the direction from the first closure spar 26 to the internal spar 62; to say in a direction orthogonal to the first closure spar 26. The widening of the duct 60 at this first region 70 makes it possible to guarantee that the air flow cross section is substantially constant within the duct 60 despite the presence of fastening members of the structural block 34 within the conduit, as will become more clearly apparent in the following. In a similar way, the duct 60 has a second region 72, which is for example located in front of the first region 70 above, and which also has an enlarged cross section in the direction from the first closure beam 26 towards the inner spar 62. The widening of the duct 60 at the level of this second region 70 makes it possible to guarantee that the effective air passage cross-section is substantially constant within the duct 60, despite the presence of a supply duct. air 74 intended to be connected to the output of the aircraft cell and input to a heat exchanger which equips the engine, in particular to perform the functions of conditioning and pressurizing the air within the cabin of the aircraft . As shown in FIGS. 2, 3 and 4, the thermal protection system 58 further comprises two air inlets 76, which for example take the form of two bent tubes each connected to the duct 60 through the inner spar 62, and opening respectively through said lateral closure walls 30 of the box 24, preferably in front of a front end of the longitudinal partition wall 64.
    [0014] Alternatively, the thermal protection system 58 may comprise a different number of air inlets, for example a single air inlet. As a further variant, as shown in FIG. 5, the air inlets may take the form of orifices 78 formed respectively in the lateral closing walls 30 facing the duct 60. These orifices are preferably extended outwardly. of the box 24 by means of air sampling tubes 80. In the example illustrated, the air inlets are connected to respective ducts (not shown) which open through the fan casing 12 into the duct of the duct. flow of secondary flow of the engine. Alternatively, the air inlets can be connected to respective ducts opening through the central casing 16, into the primary flow flow channel of the engine, within a compressor of the engine. As a variant, the air inlets may be connected to a cooling air circuit which is also intended to supply a heat exchanger whose function is to cool the air circulating in the air supply duct 74 intended to be connected. to the aircraft cell.
    [0015] FIGS. 2 to 4 also make it possible to see a front closure rib 82 of the box 24. This rib 82 connects a front end of the first closure beam 26 to a front end of the second closure beam 28.
    [0016] FIG. 3 also makes it possible to see part of the front engine attachment 6, more precisely a mounting plate 84 belonging to this engine attachment and applied to the first closure spar 26, and fixed on the latter by means of members In the illustrated example, these fasteners allow attachment of the front engine attachment 6 together on the first closure spar 26 and on a sole 88 of the front closure rib 82. As illustrated by FIG. 6, the thermal protection system 58 further comprises two air outlets, which take for example the shape of orifices 88 adjacent in the direction Y formed in the first closure beam 26 facing the duct 60, back a rear end of the longitudinal partition wall 64. These holes are preferably extended outwardly of the box 24 by means of respective air guide tubes 90. In Figure 6 illustrating a side view of the box 24, one of the air outlets is masked by the other air outlet. In the illustrated example, the air outlets open, through the air guide tubes 90, within a rear aerodynamic fairing (not shown) of the attachment pylon 4, of the type commonly referred to as " APF "or" Aft Pylon Fairing ". Alternatively, the air outlets can lead under such a rear aerodynamic fairing, a short distance from a lower heat shield of the fairing, so as to form a film of air along the heat shield. FIG. 6 also makes it possible to see a rear closing rib 92 of the box 24, connecting a rear end of the first closure beam 26 to a rear end of the second closure beam 28. FIGS. 7 and 8 more particularly illustrate the first aforementioned region 70 of the duct 60 as well as the structural block 34 arranged facing this region 70.
    [0017] In the illustrated example, widening of the duct in the first region 70 is obtained by means of a recess 94 formed in the inner spar 62. This recess 94 constitutes a portion of the spar projecting in the direction of the outside the duct 60, that is to say in the opposite direction to the first closure spar 26.
    [0018] The structural block 34 comprises an attachment plate 96 applied to the first closure beam 26 and fixed thereto by means of fasteners such as traction bolts passing through the first closure beam 26 and the inner spar 62. these fixation members are symbolized their respective axes 98. In the example illustrated, the fasteners further pass through the longitudinal partition wall 64. The fasteners are for example divided into five longitudinal rows each comprising three fasteners. Advantageously, the fasteners are guided within spacers 100 extending within the conduit 60, generally from the first closure spar 26 to the inner spar 62. For this purpose, the spacers 100 have orifices. 101 or bores for the passage of fasteners. As shown in FIG. 8, these struts 100 are for example formed of first portions 102 integrated in the first closure spar 26 and projecting in the direction of the inner spar 62, as far as the longitudinal partition wall 64, as well as second portions 104 integrated to the inner spar 62 and projecting in the direction of the first closure spar 26, also to the longitudinal partition wall 64. Each of these spacers 100 advantageously has tapered front ends 106 and rear 108 (Figure 7) so as to optimize the flow of air around each spacer. For this purpose, it should be noted that the spacers 100 are spaced apart from each other.
    [0019] Alternatively, as illustrated in FIG. 9, widening of the duct 60 in the first region 70 can be obtained by means of a recess 94a formed in the first closure spar 26. In this case, the recess 94a constitutes a portion of the first closure beam 26 projecting in the opposite direction to the inner spar 62.
    [0020] The widening of the duct 60 in the second region 72 can also be obtained by a recess formed in the inner spar 62 and / or in the first closure spar 26. In the illustrated example, this widening of the duct 60 in the second region 72 is obtained by two recesses respectively formed in the inner spar 62 and in the first closure spar 26. Referring to Figures 7 and 10, the first closure spar 26, the inner spar 62 and the longitudinal partition wall 64 comprise preferably longitudinal ribs 110 stiffening extending within the duct 60. In addition to a stiffening function to overcome buckling problems, these longitudinal ribs 110 can improve the heat exchange between the air flowing in the duct 60 and the longitudinal members 26 and 62. The longitudinal ribs 110 thus constitute heat exchange fins. Figures 7 and 10 illustrate the assembly of these elements, for example by bolting. For this purpose, the first closure beam 26 comprises, for example, lateral end fins 112 extending orthogonally to the spar, and provided at their respective free ends with respective flanges 114 extending parallel to the spar and applied to lateral edges. 115 of the longitudinal partition wall 64. Similarly, the inner spar 62 comprises, for example, lateral end fins 116 extending orthogonally to the spar and provided at their respective free ends with respective flanges 118 extending parallel to the spider. spar and applied to the lateral edges of the longitudinal partition wall 64.
    [0021] The lateral edges of the longitudinal partition wall 64 are thus sandwiched between the flanges 114 and 118, and the assembly thus formed is secured by means of bolts, or equivalent devices, defining fixing pins 119. As illustrated, the lateral end fins 116 of the inner spar 62 also extend beyond this spar 62 in the direction of the second spar of closure of the box, and are respectively connected to side panels 120 of the box, for example to the by means of bolts defining attachment pins 121. Thus, the side walls 30 of the box are each formed of a side panel 120 (or of several adjacent panels), as well as lateral end fins 112 and 116 which extend substantially in the extension of this side panel 120. It should be noted that the longitudinal members 26, 28, 62, the closing side walls 30, and the longitudinal partition wall 64 are of preferably made of metal, for example aluminum, titanium, steel, or titanium-based alloy. These elements can be made by machining in the mass, or by means of parts assembled by welding and possibly remanufactured. FIG. 11 illustrates an alternative embodiment in which the assembly of the elements described above is carried out by means of fishplates 122, each in the form of a plate. Each splint 122 includes a first portion 124 fixed to a corresponding side panel 120 of the box, and a second portion 126 extending beyond the side panel 120. The side end fins 112 of the first closure spar 26 are devoid of flanges and are directly applied and fixed to the respective second portions 126 of the fishplates 122. It is the same with respect to the end lateral fins 128 of the longitudinal partition wall 64.
    [0022] The lateral end fins 116 of the inner spar 62 are directly applied and fixed to the first respective portions 124 of the fishplates 122. The fishplates 122 may be planar. Alternatively, when the side panels 120 of the box have a slight inclination as in the example shown, the ribs 122 may form a slight angle between their respective first and second portions 124, 126. In operation, the relatively fresh air taken by means of the air inlets 76 circulates within the duct 60 and leaves the duct 60 through the air outlets 90. The circulation of the air within the duct 60 makes it possible to protect the box 24 of the rigid structure 10 of the attachment pylon 4 with respect to the heat radiated by the turbojet engine core.
    [0023] The thermal protection system also makes it possible to fulfill a firewall function in the event of an engine fire, so as to contain the fire in a dedicated area called "fire zone". In order to further improve the efficiency of the thermal protection system 58, it may further comprise a thermal protection mat disposed under the first closure spar 26, that is to say on the face of the latter located on the side 24. This type of thermal protection mattress consists of insulating materials (microporous type, airgel, etc.) and increases the efficiency of the thermal protection system.
    [0024] Other assembly modes of the elements described above are of course possible without departing from the scope of the invention.
    权利要求:
    Claims (15)
    [0001]
    REVENDICATIONS1. A propulsion unit (1) for an aircraft, comprising a motor (2) and an attachment pylon (4) intended for attaching the engine to the wing (3) of an aircraft, said attachment pylon comprising: - a rigid structure (10) comprising a box (24) comprising a first closure spar (26), a second closure spar (28) opposite the first closure spar, and two lateral closure walls (30) each having a first end attached to the first closure beam (26) and a second end attached to the second closure beam (28); - a mounting system (11) for fastening the motor (2) to said box (24), and - a thermal protection system (58) comprising a duct (60) forming a thermal barrier to protect said box (24) of the engine heat (2), characterized in that said duct (60) is housed inside said box (24).
    [0002]
    2. propulsion unit according to claim 1, wherein: - said box (24) comprises an inner spar (62) having lateral ends connected to said closing side walls (30), said inner spar (62) extending between said first closure beam (26) and said second closure beam (28); and said duct (60) is delimited by the first closure spar (26), the inner spar (62) and the lateral closing walls (30).
    [0003]
    3. The propulsion assembly according to claim 2, wherein said mounting system (11) comprises a rear engine attachment (8), and said rigid structure (10) of the attachment pylon (4) comprises a structural block (34) comprising a fixing plate (96) applied to said first closure spar (26) and fastened to the latter by means of fasteners (98) passing through the first closure spar (26) and said inner spar (62); , and- an interface for fixing (36) said rear engine attachment (8).
    [0004]
    The propulsion assembly according to claim 3, wherein said thermal protection system (58) has spacers (100) connecting said first closure spar (26) to said inner spar (62) through said conduit (60), each spacer (100) integrating at least one orifice (101) for the passage of one of said fasteners (98) of said structural block.
    [0005]
    5. A propulsion assembly according to claim 3 or 4, wherein a first region (70) of said duct (60) facing said structural block (34) has an enlarged transverse section in the direction from said first closure spar ( 26) to said inner spar (62).
    [0006]
    6. propulsion unit according to any one of claims 2 to 5, wherein a second region (72) of said duct (60), through which an air supply duct (74) intended to be connected to the air duct an aircraft has an enlarged cross section in the direction from said first closure beam (26) to said inner spar (62).
    [0007]
    A propulsion assembly according to any one of the preceding claims, wherein said mounting system (11) comprises a forward engine attachment (6) having a mounting plate (84) applied to said first closure beam (26) and secured on the latter.
    [0008]
    A propulsion assembly according to any one of the preceding claims, wherein said thermal protection system (58) has a longitudinal partition wall (64) extending between said first closure spar (26) and said inner spar ( 62) and which is connected to said closing side walls (30) so as to share a middle region of the two-part conduit (66a, 66b).
    [0009]
    A propulsion assembly according to any one of the preceding claims, wherein said box (24) comprises a front closure rib (82) connecting a front end of said first closure beam (26) to a front end of said second closure spar. (28).
    [0010]
    A propulsion assembly according to any one of the preceding claims, wherein said box (24) comprises a rear closure rib (92) connecting a rear end of said first closure beam (26) to a rear end of said second closure spar. (28).
    [0011]
    A propulsion assembly according to claim 2 or any one of claims 3 to 9 taken in combination with claim 2, wherein said box (24) has internal stiffening ribs (32) having an end connected to said inner spar. (62) and an opposite end connected to said second closure beam (28).
    [0012]
    12. propulsion unit according to any preceding claim, wherein said thermal protection system (58) comprises at least one air inlet (76, 78) connected to said conduit (60) and opening through one of said closing side walls (30) or through said first closure beam (26).
    [0013]
    13. Propulsion unit according to any one of the preceding claims, wherein said thermal protection system (58) comprises at least one air outlet (88) connected to said conduit (60) and opening through one of said walls. closing side (30) or through said first closure beam (26).
    [0014]
    A propulsion assembly according to any one of the preceding claims, wherein said thermal protection system (58) comprises longitudinal stiffening lines (110) against buckling, said longitudinal ribs (110) forming heat exchange fins.
    [0015]
    15. Aircraft, characterized in that it comprises at least one propulsion unit (1) according to any one of the preceding claims.
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    FR2960519A1|2011-12-02|Aerodynamic fairing i.e. rear lower aerodynamic fairing, for hooking device i.e. hooking strut, of turbo-jet engine in aircraft, has stiffener including pressed flange extending along stiffener direction
    FR3000721A1|2014-07-11|AIRCRAFT PROPULSIVE ASSEMBLY COMPRISING AERODYNAMIC AERODYNAMIC REAR FITTING FOR SIDE WALLS FOR THE INJECTION OF FRESH AIR ALONG A THERMAL PROTECTION FLOOR
    FR3011584A1|2015-04-10|EXTENSION OF INTERMEDIATE CASING
    CA2786542A1|2011-09-09|Turbojet engine nacelle provided with a cooling assembly for cooling a component
    EP3581781B1|2021-07-28|Propulsion system of an aircraft comprising a fixed interior structure having an evacuation slot
    FR3053402A1|2018-01-05|ARRANGEMENT FOR THE FASTENING OF A THRUST INVERSION TRAP ROD ON A FIXED INTERNAL STRUCTURE OF A TURBOJET NACELLE AND ASSOCIATED MOUNTING / DISMANTLING METHOD
    FR2964947A1|2012-03-23|Lower aft pylon fairing for engine mounting structure inserted between wing and turbojet engine of aircraft, has extension traversing inner cross stiffening rib and fixed on spar, where extension and bracket are formed as single piece
    同族专利:
    公开号 | 公开日
    CA2890421A1|2015-11-09|
    FR3020798B1|2017-12-22|
    US20150321765A1|2015-11-12|
    US9975641B2|2018-05-22|
    CN105083565A|2015-11-25|
    CN105083565B|2019-07-30|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US3893638A|1974-02-14|1975-07-08|Boeing Co|Dual cycle fan jet engine for stol aircraft with augmentor wings|
    US5123242A|1990-07-30|1992-06-23|General Electric Company|Precooling heat exchange arrangement integral with mounting structure fairing of gas turbine engine|
    FR2891255A1|2006-07-07|2007-03-30|Airbus France Sas|Engine e.g. jet engine, assembly for aircraft, has outlet placed in rear with respect to rear engine mount, and heat exchanger system with exchanger arranged inside fairing that is entirely situated in rear with respect to engine mount|
    FR2931133A1|2008-05-14|2009-11-20|Airbus France Sas|MOTOR ATTACHING MACHINE COMPRISING MEANS FOR FIXING LONGERONS AND AGENCY PANELS OUTSIDE OF THE INTERIOR SPACE OF HOUSING|
    US20120080554A1|2010-10-01|2012-04-05|Airbus Operations |Aircraft engine attachment pylon comprising two front wing system attachments with orthogonal shearing pins|
    US4712750A|1986-05-02|1987-12-15|The Boeing Company|Temperature control device for jet engine nacelle associated structure|
    FR2891250B1|2005-09-28|2007-10-26|Airbus France Sas|AIRCRAFT ENGINE ASSEMBLY COMPRISING AN ENGINE AND A HITCHING MACHINE OF SUCH AN ENGINE|
    FR2891248B1|2005-09-28|2009-05-01|Airbus France Sas|ENGINE ASSEMBLY FOR AN AIRCRAFT COMPRISING AN ENGINE AND A MACHINE FOR ATTACHING SUCH A MOTOR|
    FR2921342B1|2007-09-20|2010-03-12|Airbus France|LOWER REAR AERODYNAMIC FAIRING FOR AN AIRCRAFT ENGINE CLAMPING DEVICE|
    US7943227B2|2007-10-11|2011-05-17|The Boeing Company|Ceramic heat shield|
    FR2960522B1|2010-05-27|2012-06-29|Airbus Operations Sas|METHOD FOR MANUFACTURING BY SUPERPLASTIC FORMING AND BY LAUNDRYING A RIB FOR AERODYNAMIC FITTING OF AN AIRCRAFT ENGINE HITCHING MAT|
    GB201011056D0|2010-07-01|2010-08-18|Rolls Royce Plc|Pylon for attaching a gas turbine engine|
    FR2977237B1|2011-06-28|2014-11-21|Airbus Operations Sas|REAR AERODYNAMIC FAIRING OF A MATERIAL FOR CONNECTING AN AIRCRAFT ENGINE|
    US9238511B2|2014-03-04|2016-01-19|Mra Systems, Inc.|Engine pylon structure|FR3009339B1|2013-07-30|2018-01-26|Safran Aircraft Engines|TURBOMACHINE COMPRISING A PYLON COOLING DEVICE|
    FR3032180B1|2015-01-30|2018-05-18|Airbus Operations|PROPELLANT ASSEMBLY COMPRISING A TURBOJET ENGINE AND A COUPLING MAT FOR A NEW DISTRIBUTION OF EFFORTS BETWEEN THE TURBOJET AND THE VIL|
    US10907500B2|2015-02-06|2021-02-02|Raytheon Technologies Corporation|Heat exchanger system with spatially varied additively manufactured heat transfer surfaces|
    US10487744B2|2016-05-23|2019-11-26|United Technologies Corporation|Fence for duct tone mitigation|
    US11156165B2|2019-06-13|2021-10-26|The Boeing Company|Fire seal assemblies for aircraft engines|
    法律状态:
    2015-05-21| PLFP| Fee payment|Year of fee payment: 2 |
    2015-11-13| PLSC| Publication of the preliminary search report|Effective date: 20151113 |
    2016-05-20| PLFP| Fee payment|Year of fee payment: 3 |
    2017-05-23| PLFP| Fee payment|Year of fee payment: 4 |
    2018-05-22| PLFP| Fee payment|Year of fee payment: 5 |
    2019-05-22| PLFP| Fee payment|Year of fee payment: 6 |
    2020-05-22| PLFP| Fee payment|Year of fee payment: 7 |
    2022-02-11| ST| Notification of lapse|Effective date: 20220105 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1454183A|FR3020798B1|2014-05-09|2014-05-09|AIRCRAFT PROPULSIVE ASSEMBLY COMPRISING A THERMAL BARRIER CONDUIT INTEGRATED WITH THE HOUSING OF THE RIGID STRUCTURE OF THE ATTACHING MAT|FR1454183A| FR3020798B1|2014-05-09|2014-05-09|AIRCRAFT PROPULSIVE ASSEMBLY COMPRISING A THERMAL BARRIER CONDUIT INTEGRATED WITH THE HOUSING OF THE RIGID STRUCTURE OF THE ATTACHING MAT|
    CA2890421A| CA2890421A1|2014-05-09|2015-04-29|Aircraft propelling assembly including a duct forming a thermal barrier integrated in the caisson of the rigid structure of the engine mounting system|
    CN201510232377.4A| CN105083565B|2014-05-09|2015-05-08|Propulsion assembly and aircraft for aircraft|
    US14/707,240| US9975641B2|2014-05-09|2015-05-08|Aircraft propelling assembly including a duct forming a thermal barrier integrated in the caisson of the rigid structure of the engine mounting system|
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