Aircraft engine assembly comprising an annular load-transfer structure surrounding the central casing of a turbojet engine

Translation of description of equivalent WO 2009112781 A2

AIRCRAFT ENGINE ASSEMBLY INCLUDING A STRUCTURE

RING TRANSFER EFFORT SURROUNDING THE CARTER

CENTRAL OF A JET ENGINE

DESCRIPTION

p0005The present invention relates generally to an aircraft engine assembly of the type comprising a turbojet, a nacelle surrounding the jet engine, and an engine mount provided with a rigid structure and a plurality of engine attachments interposed between a rigid structure of the suspension pylon and the turbojet.

p0006The suspension pylon, also called "EMS" ( "Engine Mounting Structure") for suspending the turbojet engine below the wing of the aircraft, or to mount this turbojet above this same wing, or even to bring it back fuselage. It is designed to form the connecting interface between a turbojet engine and a particular structural part of the aircraft. It transmits to the structure of this aircraft the forces generated by its associated turbojet, and also enables routing of fuel, electrical, hydraulic, and air between the engine and 1 aircraft. The pod is itself conventionally equipped with several covers enveloping turbojet and allowing access to the latter in open position, these covers being known under the names of fan cowl and thrust reverser cowls. More specifically, on some motor assemblies of the prior art, there is provided a mounting pylon having a rigid structure comprising a box longitudina l and two lateral caissons integral with the longitudinal housing and arranged on either of other thereof, the mast further comprising catching means of the jet engine on the rigid structure, these means comprising a first, second and third forward engine attachments of recovery of the thrust forces to the fan housing. As schematically shown in Figure 1 illustrating an embodiment of the prior art in which the motor is intended to be suspended under the wing of the aircraft, three forward engine attachments containing the thrusts are arranged so that the third front engine 8 through a diametral plane Pl of the turbojet, here the vertical symmetry plane of the turbojet engine, while the first and second engine 6a, 6b, respectively intended to be reported on the two lateral panels of the mast, are about they arranged on either side of this diametral plane Pl, and usually traversed by another diametrical plane P2 of the turbojet, orthogonal to the said diametral plane and corresponding here to the horizontal plane of symmetry of the turbojet. Furthermore, the jet engine conventionally comprises a fan casing 12, an intermediate casing 21 located radially inwardly relative to the fan casing and connected thereto by a plurality of structural arms 17 oriented preferably radially, as well as a central casing 16, also housing said "core", extending the intermediate casing 21 rearwardly. Finally, it is noted that the central housing extends to a rear end 19 of larger size, also called casing ejection.

p0007As shown in Figure 2, the engine assembly includes an annular structure of transfer efforts 60 surrounding the central casing 16 and mechanically connected thereto via mounting means 62, typically comprising a plurality of connecting rods . As shown diagrammatically by the arrows 76, the annular structure 60 is also connected to a plurality of structures (not shown) externally arranged with respect thereto, and urging the example radially, respectively in a plurality of points introduction of effort, generally circumferentially distributed on the latter. Thus, the annular structure allows to pass the forces between the central casing and the field, these being for example the delimitation of outer radial structure of the annular secondary flow channel (from the English "OFS" Fan Outlet Structure ) and / or the radially inner boundary structure of the channel annular secondary flow (from the English "IFS" Inlet Fan Structure).

p0008However, in the solutions of the prior art, the provision of the aforementioned connecting rods are not optimized, the passage of forces causes deformations of the ring structure and / or of the central casing, which is of course not desirable.

p0009The invention therefore aims to propose an aircraft engine assembly at least partially overcomes the above mentioned problems related to embodiments of the prior art.

p0010To do this, the invention relates to an aircraft engine assembly comprising: - a turbojet engine comprising a fan casing, an intermediate casing located radially inwardly relative to the fan casing and connected thereto by a plurality of structural arms, preferably oriented radially, and a center housing extending said intermediate casing rearwardly; an annular structure surrounding effort transfer the center housing and mechanically connected thereto by means of mounting means including a plurality of connecting rods, said annular structure being also connected to a plurality of substantially planar structures arranged externally with respect thereto, and urging the respectively in a plurality of force introduction points distributed circumferentially on the latter. According to the invention, at least one connecting rod is associated with each insertion effort point, said link being, in front view according to a longitudinal axis of the turbojet, disposed tangentially to the central casing and having a inner end connected to the central casing and an outer end connected to said annular structure so as to be traversed by an imaginary plane in which lies said structure, and passing through said point of introduction effort.

p0011The invention has the advantage of placing, in front view, each point of introduction of effort and the outer end of the connecting rod in a same imaginary plane, preferably radial, wherein is also arranged the effort introduced into the annular structure, the outer structure of the same plane associated with the point of introduction of forces concerned. Therefore, the aforementioned load, substantially radial or not, is jointly transmitted by a compression or tensile stresses in the connecting rod, and by a substantially tangential force necessarily in the ring structure, also said membrane effort. Therefore, at each of the points of introduction of efforts of the annular structure, it tends to respond to mechanical stress field structure by a substantially tangential force, greatly limiting its deformations, and especially preventing to "become oval." Furthermore, the connecting rods being disposed tangentially to the central housing on which they are connected, the deformations of this casing are also severely limited.

p0012The invention therefore provides generally provide better optimized mechanical junction between the annular structure and the central casing, to ensure a satisfactory recovery efforts transiting through the external structures connected to the ring structure.

p0013Preferably, as mentioned above, said structures are arranged substantially radially and substantially radially urge said annular structure. Each notional plane is substantially radial, passing through the longitudinal axis of the turbojet. More generally, each structure preferably extends in an imaginary plane substantially parallel to the longitudinal axis of the turbojet, or integrating the same axis.

p0014Preferably, said connecting rods are arranged substantially in the same transverse plane of the turbojet.

p0015Still preferably, in order to permit the better the thermal expansion of the center housing with respect to the annular structure that surrounds said connecting rods have inner and outer ends mounted ball joint manner. Preferably, said connecting rods all extend in the same direction circumferential from their outer end. With this configuration, in case of differential thermal expansion between the ring and the housing or the connecting rods, the ring may advantageously rotate about the central casing, while remaining coaxial thereto.

p0016Preferably, the engine assembly further comprises a mounting pylon having a rigid structure and hooking means of said turbojet engine on the rigid structure, said mounting means comprising a first, second and third forward engine attachments of recovery of the thrust forces to the fan housing, and arranged so that said third forward engine attachment passes through a first diametral plane of the turbojet engine, said first and second forward engine attachments are arranged on either side of this first plane diametral, and in that each of said first, second and third forward engine attachments is associated a reinforcing structure forming a shear plane formed by one of said aforementioned structures and fixedly connected:

p0017- At the level of the annular structure into a first anchoring point forming said force introduction point in said annular structure;

p0018- At the level of the fan casing in a second anchor point; and at one or structural arm of the intermediate casing in a third anchor point, said reinforcing structure extending along said imaginary plane passing through said introduction point of effort, and also by an anchor point of said forward engine attachment on the fan case.

p0019The presence of these shear stressed reinforcing structures serves to stiffen the turbojet in these three imaginary planes, implying a limitation of the bending of the central casing and the intermediate casing, even if the inertial stresses in these areas. The overall performance of the engine are improved.

p0020Furthermore, this addition of reinforcing structures provides structural stiffening arm and close to the two imaginary planes associated with first and second brackets, that is to say where the arms are traditionally the most requested. This advantageously results in lower deformation of the structural arms. Therefore, the fan casing is less likely to open up in terms of structural arms, which greatly limits the effect of roundness met with first and second diametrically opposed engine mounts. This results in better performance of the fan, and thus better overall efficiency of the engine.

p0021In addition, the shear plane at reinforcing structures fully play their role through non-deformation mentioned above of the annular structure on which they are connected, through the points of introduction effort.

p0022Still preferentially, said reinforcing structures are devoid of direct mechanical connection with said mast, which avoids introducing additional efforts in the latter. The aforesaid hooking means can thus remain isostatic, despite the presence of the reinforcing structures. As an indication, for identical reasons, it also ensures that said reinforcing structures are devoid of direct mechanical connection with the nacelle of the motor concerned.

p0023Preferably, said first and second forward engine attachments are attached to the fan casing respectively at two points situated beyond the second diametral plane of the turbojet orthogonal to the first diametral plane relative to said third forward attachment motor. Alternatively, the first and second forward engine attachments could be to the fan housing at two points respectively located in said second diametrical plane, without departing from the scope of the invention.

p0024Preferably, said first and second engine before resumption of the thrust forces are located symmetrically with respect to said first diametrical plane defined by the longitudinal axis of the turbojet parallel to a longitudinal direction thereof, and a first direction of said turbojet orthogonal to the longitudinal direction. Still preferably, the first and second front engine mounts are each designed so as to resist forces applied along the longitudinal direction and along the first direction of the turbojet, and said third front attachment engine is designed to transfer loads' exerted in the longitudinal direction and in a second direction of the turbojet, orthogonal to said first direction and in the longitudinal direction.

p0025As an indication, it is noted that where the jet engine is intended to be mounted above the wing of the aircraft or suspended below it, the first and second directions orthogonal to each other and orthogonal to the longitudinal direction are preferably the vertical and transverse directions of the turbojet, respectively. However, although this may also be the case in connection with the attachment of the motor assembly in rear part of the fuselage of the aircraft, it may be that the first and second directions are each inclined with respect to vertical directions and cross the turbojet.

p0026In this configuration, said fastening means are only constituted by the front attachments above, attached to the fan casing of the turbojet, and forming an isostatic recovery system. In a more general way, it is ensured that the only attachment means secured to the fan casing are said first, second and third engine mounts, even in other cases where a fastener additional motor is provided between the rigid structure of the strut and the central casing, always in order to form an isostatic recovery system.

p0027Preferably, as mentioned above, said first direction of the turbojet corresponds to a vertical direction thereof, and said second direction of the turbojet corresponds to a transverse direction thereof.

p0028Another object of the present invention concerns an aircraft comprising at least one engine assembly as described above, assembled on a wing or on a rear fuselage portion of this aircraft.

p0029Other advantages and features of the invention appear in the detailed non-limiting description below.

p0030This description will be made with reference to the accompanying drawings in which; Figures 1 and 2, already described, show an engine assembly for aircraft according to the prior art;

p0031- Figure 3 shows a side view of an engine assembly for an aircraft, according to a preferred embodiment of the present invention; 4 shows a perspective view of the assembly shown in Figure 3, the rigid structure of the strut, the reinforcing structures and the nacelle having been removed to allow more clearly show the engine mounts; 5 shows a schematic front view corresponding to that of Figure 4, illustrating the particular positioning of the engine mounts;

p0032- Figure 6 shows a partial enlarged perspective view of the engine mount according to the preferred embodiment view;

p0033- Figure 7 shows a sectional view along the transverse plane P 'of Figure 6; 8 shows a perspective view of an attachment strut part of an engine assembly for an aircraft according to another preferred embodiment of the present invention;

p0034- Figure 9 shows a sectional view along the transverse plane P 'of Figure 8, through the rigid structure of the pylon; - Figure 10 shows a cross-sectional view illustrating the reinforcing structures connecting the fan casing to the central housing, this figure also corresponds to a sectional view taken along the line XX of Figure 11; - Figure 11 shows a sectional view taken along the line XI-XI of Figure 10;

p0035- Figure 12 shows a sectional view taken along the line XII-XII of Figure 10;

p0036- Figure 13 shows a cross-sectional view diagrammatically showing the mounting means between the annular structure reported effort transfer, and the central casing of the turbojet engine; Figure 13a shows a similar view to that of Figure 13, the mounting means being under an alternative embodiment; - Figure 14 shows a partial and enlarged view of that shown in Figure 13, showing diagrammatically the distribution of forces on the annular structure at a point of introduction of effort; and

p0037- Figure 15 shows another preferred embodiment, corresponding to a sectional view taken along the line XV-XV of Figure 13.

p0038Referring to FIG 3 shows an engine assembly 1 for an aircraft according to a preferred embodiment of the present invention, this assembly 1 being intended to be fixed under an aircraft wing (not shown).

p0039Overall, the motor assembly 1, also called integrated propulsion system consists of a turbojet engine 2, a nacelle 3 (shown in phantom lines for clarity reasons), and a mount 4 provided with means hooking of the turbojet on the mast, these means being preferably constituted by a plurality of engine attachments 6a, 6b, 8, reported fixedly on a rigid structure 10 of the mounting pylon (attachment 6b being masked by the mount 6a in FIG 3). As an indication, it is noted that the assembly 1 comprises another series of attachments (not shown) to ensure the suspension of this assembly 1 under the wing of the aircraft.

p0040Throughout the following description, by convention, X refers to the longitudinal direction of the pylon 4 which is also comparable to the longitudinal direction of the turbojet 2, this direction X being parallel to a longitudinal axis 5 of this turbojet 2. Furthermore, Y designates the direction oriented transversely to the mast 4 and also comparable to the transverse direction of the turbojet 2, and Z the vertical direction or height, these three directions X, Y and Z orthogonal to each other.

p0041Furthermore, the terms "front" and "rear" are to be considered with respect to a direction of travel of the aircraft subsequent to the thrust exerted by the turbojet 2, this direction being schematically represented by arrow 7.

p0042In Figure 3 it can be seen that only the engine attachments 6a, 6b, 8 and the rigid structure 10 of the mount 4 are shown. The other constituents not shown of this pylon 4, such as the attachment means 10 of the rigid structure under the wing of the aircraft, or the secondary structure segregating and maintaining while supporting aerodynamic fairings are conventional elements identical or similar to those encountered in prior art and known to the skilled person. Therefore, it can be made therewith no detailed description. On the other hand, the turbojet engine 2 has an identical or similar design to that shown in Figure 1, namely comprising at the front a fan casing 12 of large size delimiting an annular fan duct 14, an intermediate casing 21 and structural arms 17 (not shown in Figure 3), also called blades output managers, and a central housing 16 having a rear end 19.

p0043As is apparent from the above, this is preferably a jet engine with a high bypass ratio.

p0044As can be seen in FIG 3, a first forward engine attachment 6a and a second forward engine attachment 6b are both intended to be fixed on the fan housing 12, symmetrically relative to a plane P said first diametral plane defined by the axis 5 and the Z direction, this vertical plane Pl passing through a third forward engine attachment 8 also fixed to the fan casing 12, the three fasteners being preferably all traversed by a plane orthogonal to one axis 5.

p0045Referring now to Figure 4, one can see that the first attachment 6a and the second mount 6b schematically represented are actually arranged symmetrically relative to the first diametral plane Pl of the turbojet, and preferably both arranged on a part annular peripheral of the fan casing 12, specifically on the back of this same part. In this preferred embodiment, they are arranged in a plane P2 said second diametrical plane of the jet engine, which is orthogonal to the first, and therefore horizontal. The two connection points 6'a and 6'b for fastening means 6a, 6b on the housing 12 are therefore located so that the second plane P2 is disposed between on the one hand these two points 6'a and 6 ' b, and secondly, a connection point 8 'of the engine attachment 8 on the same casing, in front view along the axis 5, as that of Figure 5.

p0046In this figure a view along the axis 5, it can be seen that Al angle having as center the longitudinal axis 5, between the attachment points 8 'and 6' has the third and the first engine attachment is strictly greater than 90 ° and preferably between 90 and 110 ° not included. Similarly, an angle A2 having as center the longitudinal axis 5, between the attachment points 8 'and 6'b of the third and of the second engine attachment is strictly less than 270 °, and preferably between 250 270 ° non included. This arrangement of the attachments 6a, 6b allows solicit more engine attachment 8, and therefore limit the parasitic effects of roundness of the fan casing encountered in embodiments of the prior art, with the first and second engine fasteners arranged in the plane P2. In addition, it helps counter / offset a torque axis parallel to the Y direction, exerted on the jet engine, resulting axial forces passing by that third fastener 8. Anyway, although this provision is preferred, the invention also applies to other provisions of the engine mounts 6a, 6b, 8.

p0047As an indication, it is noted that the engine attachments 6a, 6b, 8 are made conventionally, for example of integrating hardware and axes, the anchors / de aforementioned connection 6'a, 6'b, 8 'corresponding to the points of contact between the structure of these ties and the fan casing structure.

p0048As shown schematically by the arrows in Figure 4, each of the first and second forward engine attachments 6a, 6b is designed so as to resist forces generated by the turbojet engine 2 in the X direction and the Z direction, but not those exerted in the direction Y.

p0049In this way, the two attachments 6a, 6b away from each other jointly assure the transfer of the moment exerted in the X direction, and that the moment exerted in direction Z. Still referring to Figure 4, one can see that the third front attachment 8 located on the highest part of the fan casing 12, thus the uppermost portion of the peripheral annular part is designed so as to resist forces generated by the turbojet engine 2 in the X direction and in the direction Y, but not those exerted in direction Z. In this way, the third mount 8 assures jointly with attachments 6a, 6b absorbing the torque exerted in the direction Y .

p0050The advantage of this non-limiting configuration lies in the fact that all the engine mounts are mounted on the fan casing, so that the secondary stream is in no way disturbed by these fasteners, resulting in a significant gain in terms of overall performance engine. Furthermore, three ties together an isostatic recovery system.

p0051Referring now to FIG 6 shows an embodiment of the rigid structure 10 of the engine mount 4. First, it is indicated that this rigid structure 10, also called the primary structure, is preferably designed so as to be symmetrical to the diametral plane P indicated above, that is to say relative to the vertical plane defined by the longitudinal axis 5 of the turbojet 2, and the Z direction indication, it is generally the case when the engine is suspended or mounted above the wing, but not necessarily met when it is assembled to the rear fuselage. Indeed, in the latter case will be detailed with reference to Figures 8 and 9, the rigid structure 10 may have a plane of symmetry another depending on its orientation relative to the rear fuselage, for example a substantially horizontal symmetry plane or inclined relative to the horizontal, or even present no plane of symmetry. This occurs in particular when the two lateral caissons which will be described hereinafter, integral and disposed on either side of a longitudinal box said central box, do not have the same circumferential length.

p0052Thus, the rigid structure 10 has a longitudinal box 22, said longitudinal central box, and also called torsion box, which extends from one end to the other of the structure 10 in the X direction, parallel to the same direction . AT indication, this box 22 may be formed by assembling two side members or side panels 30 extending in the X direction in parallel planes XZ, and connected together via transverse ribs 25 which are for their oriented in parallel YZ planes. In addition, an upper spar 35 and a lower beam 36 are also provided for closing the box 22.

p0053Two lateral caissons 24a, 24b complete the rigid structure 10, the central box 22 is located at an upper portion of the same structure 10, each of the two chambers 24a, 24b being integral with the central torsion box 22 and projecting to either side thereof along the Y direction, and down. As an indication, it is noted that the wells 22, 24a, 24b could be constructed so as to form a single box, within the scope of one invention. Preferably, these side casings reported integrally to both sides at the front of the central box 22, each have an inner skin 26a of box closure, 26b, also called the lower skin, oriented turbojet and jointly defining a portion a fictitious cylindrical surface 32 of substantially circular section and a longitudinal axis 34 parallel to the central box 22 and to the X direction, as can be seen in Figure 6. in other words, these two skins 26a, 26b each have at least a portion with a curvature adapted to be positioned around and in contact with said imaginary surface 32. It is advantageously provided that the skins 26a, 26b then participate in the radially outer delimitation from an annular secondary flow channel (not shown), knowing that it is still possible to provide an acoustic protective coating on these closing skins, either on their inner or outer surfaces. Alternatively, it is possible to ensure that the side casings are wholly located above the fan case, without departing from the scope of the invention.

p0054As an indication, it is specified that the axis 34 is preferably coincident with the longitudinal axis 5 of the turboprop 2.

p0055Moreover, the lateral notch 24a, here identical and symmetrical to the lateral notch 24b has an outer skin 44a box closure, while the lateral box 24a includes its own outer skin 44b of the box closure.

p0056These outer skins 44a closures, 44b, also called upper skin, are preferably each a portion of the outer aerodynamic surface of the nacelle, preferably involving at least part of the mast is an integral part of the nacelle.

p00577 shows a sectional view according to a plane P 'passing through transverse in any way the lateral caissons 24a, 24b. In this figure, one can actually see that the two inner skins closure box 26a, 26b define a portion of their outer surface a part of the imaginary surface 32 substantially circular-cylindrical. It is noted that to create the least possible disturbance of the secondary flow escaping from the annular fan duct 14, the diameter of the cylindrical imaginary surface 32 is preferably substantially identical to the diameter of the cylindrical outer surface of the annular portion of the housing blower 12. this specificity is well understood in the sense of it to provide that the skins 26a, 26b participate in the radially outer boundary of this annular secondary flow channel.

p0058On the other hand, as can be seen in Figure 7, the elements of the central box 22 will protrude over a very small distance within the space 38 delimited by the imaginary surface 32, so that they do not significantly disrupt the flow of the secondary air stream. This is explained by the fact that the side rails 30 have a height in the Z direction that is extremely small compared to the diameter of fictitious external surfaces 32 and 18.

p0059Referring jointly to Figures 6 and 7, the skins 26a, 44a are connected to each other via a forward closure frame 28a and a rear closure frame 46a, the frames 28a, 46a thus being oriented transversely and located respectively at the front and rear of the housing 24a. In addition, a cover plate 48a located below the plane P2 closes a lower portion of box 24a, and thus connects the lower end of the frames 28a, 46a and skins 26a, 44a.

p0060Naturally, the lateral box 24b comprises 26b elements 44b, 28b, 46b and 48b, respectively identical to the elements 26a, 44a, 28a, 46a and 48a 24a chamber, these two chambers being for example may preferably carry so articulated hoods of the nacelle.

p0061Both skins 26a, 26b are preferably made in one piece and interconnected at their upper part by means of a junction plate 50 oriented in an XY plane, and located in contact with the lower spar 36 the central box 22. Similarly, it is also possible that the two closure frames before 28a, 28b are formed integrally and connected together at their upper part by means of a closure frame 31 before the housing 22, the framework 31 being oriented along a YZ plane. Therefore, in this configuration, the frames 28a, 28b, 31 made in one piece are arranged in a plane YZ, and constitute a front end of the rigid structure 10 of the pylon 4.

p0062Thus, the rigid structure 10 of the mount 4 is fully adapted to support the engine attachments 6a, 6b, 8, since they can be easily fixed to the cross-piece made in one piece integrating the frames 28a, 28b and 31, as shown in Figure 1, and having for example a generally U-shaped, as are all of the rigid structure in front view. An alternative solution could be envisaged, in which the side boxes form a semi-cylindrical barrel and not U, additional structural elements then being provided in these boxes to deport the first and second engine mounts in the second diametrical plane P2 . This configuration is particularly advantageous in the context of an implementation of the turbojet on the mast by a vertical movement, from below.

p0063As an indicative example, all of the components of the rigid structure 10 which has just been described is carried out using metallic material, such as steel, aluminum, titanium, or to using composite materials, preferably carbon.

p0064It is recalled that if the side boxes 24a, 24b may actually have a different circumferential length, mainly in the case of hanging the whole back part of the fuselage, it is also clear that in the latter case, these wells could be reported to the central box 22 at a position other than forward part thereof, without departing from the scope of the invention. In this regard, referring now to Figures 8 and 9 shows the rigid structure 10 of an engine mount belonging to a motor assembly according to another preferred embodiment of the present invention, whose characteristic is of be intended to bring back part of a fuselage 1 of 80 aircraft. This rigid structure 10 has a substantially identical design to that described in the embodiment described above, as shown by the reference numerals corresponding to like elements or similar to those previously described.

p0065Can be seen that the main difference resulting from the coupling in the rear part of the fuselage 80, lies in the inclination of the rigid structure 10, in so far as the two lateral housings 24a, 24b together now form a portion of an envelope / substantially cylindrical cage which is no longer located about an upper half diameter, but arranged around a substantially half lateral diameter of this same turbojet engine (not shown).

p0066More specifically, the rigid structure 10 is preferably designed to be symmetrical with respect to the diametral plane Pl which is no longer vertical, but defined by the longitudinal axis 5 of the jet engine 2 and a first Z-direction orthogonal to the X direction, the first direction Z 'being inclined to Z and Y directions above, respectively corresponding to the vertical and transverse directions of the turbojet. Preferably, this plane P can be such that it rises by departing the fuselage 80, for example an angle between about 10 ° and 60 ° relative to the horizontal, that is to say by respect to any XY plane. The first forward engine attachment 6a and the second forward engine attachment 6b are both intended to be fixed to the fan casing, symmetrically relative to the plane P defined above, as shown in Figure 8. It is then provided the first and second forward engine attachments 6a, 6b are arranged beyond the diametrical plane P2 orthogonal to Pl, vis-à-vis the clip 8. again, it boils down that the diametrical plane P2 between on the one hand the two fasteners 6a, 6b, and on the other hand the engine attachment 8.

p0067Here, the plane P2 is defined by the longitudinal axis 5 and a second direction Y orthogonal to the X direction and the first direction Z ', so it is also inclined relative to the Z and Y directions

p0068As shown schematically by the arrows in Figure 8, each of the first and second forward engine attachments 6a, 6b is designed so as to resist forces generated by the turbojet engine 2 in the X direction and along the first direction Z ', but not those exerted in the direction Y '.

p0069In this way, the two attachments 6a, 6b greatly distant from the other jointly assure the transfer of the moment exerted in the X direction, and that the moment exerted in direction Z '.

p0070Still referring to Figure 8, there is shown a third forward engine attachment 8 and also shown schematically to be attached to the peripheral annular part of the fan casing (not shown), also preferably on the rear of this part. Regarding this third front attachment 8, fictitiously crossing the plane P indicated above, it is designed so that it only forces generated by the turbojet 2 along the X direction and in the direction Y ', and therefore not those exerted in the Z directions'. In this way, this third fastener

p00718 assures jointly with the other two mounts 6a, 6b absorbing the torque exerted in the second direction Y '.

p0072Finally, although this has not been shown, it is noted that it is preferably provided one or more nacelle covers mounted on the rigid structure 10, and particularly on the side boxes 24a, 24b.

p0073Figures 10 to 12, the turbojet integrates most of reinforcing structures connecting the fan casing to the central housing. In the figures, the jet engine 2 is shown in a position as set when it is suspended under the wing. Nevertheless, the described embodiment can be considered for any positioning of the turbojet, in particular when it is reported back portion of the fuselage, as shown in Figures 8 and 9.

p0074First, there is provided an annular structure 60 for transfer of effort, also known as rim or ring surrounding the central casing 16 and centered on the axis 5. This ring 60, radially spaced from the central casing 16 is mechanically connected thereto by means of mounting means 62 of the type connecting rods, as will be detailed below. Preferably, this ring 60 is located rearward of the central casing 16, for example downstream of the combustion chamber and more preferably at an inter-turbine casing in vis-a-vis a fixed component structure, ideally at the end of the high pressure turbine casing. To better support, it is preferably located to the right of a shaft bearing of the turbojet.

p0075First, there is provided a reinforcing structure forming a shear plane, associated with each of the first and second engine 6a, 6b.

p0076Thus, as regards the first engine attachment 6a, a reinforcing structure forming shear plane 64a is arranged in an imaginary radial plane 66a passing through the axis 5, and also through the anchor 6 has this mount 6a.

p0077As is best seen in Figure 11, the structure 64a is preferably in a plane substantially triangular shape, possibly pierced for a mass gain. The triangle is fixedly connected to the ring level 60 in a first anchor point 68a, at the fan casing 12, close to the point 6 'situated in the same imaginary plane 66a, a second point anchor 70a, and at the junction between a structural arm 17 and the intermediate casing 21, in a third anchor point 72a. Thus, the triangular structure 64a forming plan shear has a base parallel to and along the structural arm 17 placed in the notional plane 66a, the latter being inclined by relative to the directions Y and Z, due to the offset of the mount 6a below the diametrical plane P2.

p0078The imaginary plane 66a in which fits the triangular reinforcing structure 64a, here is radial, that it passes through the longitudinal axis 5. However, it could be arranged differently, ie parallel to the axis 5 longitudinal, without integrating. This is particularly the case where the structural arms are themselves not radial, but inclined in a transverse plane so that their axis does not intercept the longitudinal axis 5. In such a configuration, it preferably continues to to ensure that the triangular structure 64 has a base parallel to and along the structural arm 17 placed in the notional plane 66a. In other words, it is preferably provided that the triangular structure 64 is in the rearward extension of one of the structural arm 17, the arm 64a and the structure being thus situated in the same imaginary plane 66a. It is noted that this specificity is also applicable to each of the other reinforcing structures described below. Similarly, an identical or similar design is adopted for the second attachment 6b. Therefore, the figures, the numerals attached to items in relation to the reinforcing structure 64b forming shear plane arranged in a radial imaginary plane 66b, therefore bear the letter "b" instead of the letter "A" used for identical elements in connection with the reinforcement structure 64a.

p0079Thus, the structures 64a, 64b are found symmetrical with respect to the diametral plane Pl, also corresponding to another imaginary radial plane 66c, wherein there is a third structure 64c reinforcement forming shear plane attached to the third engine 8. in this regard, it is noted that in the possible case where the engine mounts 6a, 6b are arranged in the plane P2, and not below it, the two radial imaginary planes 64a, 64b would then be confused with this plane P2 .

p0080Here too, the figures, the numerals attached to items in relation to 64c reinforcing structure forming arranged shear plane in a radial imaginary plane 66c, therefore bear the letter "c" instead of the letter "a" used for identical elements in connection with the reinforcement structure 64a.

p0081The three structures 64a, 64b, 64c, preferably substantially identical, generally stiffen the central casing 16, thus limiting its bending, even when inertial loads exerted in the notional planes 66a, 66b, 66c, the corresponding final here the vertical plane. Furthermore, they allow a limited deformation of the structural arms 17 in the notional planes and close, advantageously resulting in a limitation of the effect of ovalization of the fan casing 12. Preferably, the structures 64a, 64b, 64c can each act as bifurcations of the air in the secondary flow of the turbojet, these bifurcations whose main function is to integrate the passage of systems and / or with an acoustic treatment, while component of the aerodynamic surfaces.

p0082Finally, for the preservation of a recovery isostatic efforts, the reinforcing structures are devoid of any direct mechanical connection with said mast, and also with the nacelle.

p0083Turning now to Figures 13 and 14, one can see one of the features of the present invention residing in the design of mounting means 62 reported between the annular structure 60 transfer efforts, and the central casing 16.

p0084First, it is noted that the anchor points 68a, 68b, 68c mentioned above each form a force introduction point in the ring 60, these points being distributed circumferentially along the latter. In addition, because of the radial preferred orientation of the aforementioned reinforcing structures associated with these items 68a, 68b, 68c, the force biasing the ring 60 is also oriented radially, ie via a direction crossing the axis 5 , which is centered on the same ring. However, it is recalled that the reinforcement structures could be oriented differently than radially within the scope of the invention. At each of these three points 68a, 68b, 68c is assigned at least one connecting rod 62, each connecting rod being, in front view along the axis 5 as shown in Figure 13, disposed tangentially to the central casing 16. More specifically, the links 62 are preferably all arranged substantially in the same transverse plane of the turbojet.

p0085It is preferably provided that a single rod 62 comes from every point 68c upper and lower 68a, 68b.

p0086For each of these links 62, is provided an inner end 62a connected so as ball joint on the central casing 16 and an outer end 62b connected manner on ball joint ring 60. More particularly, the outer end 62b is arranged so as to be traversed, in front view, for the imaginary radial plane 66a, 66b, 66c passing through the longitudinal axis 5 and the point of introduction of forces 68a, 68b, 68c concerned. As an indication, the result, in general, that for a given engine attachment, the anchor of the clip on the fan casing, the anchor point of the reinforcing structure associated to the casing fan, the reinforcing structure itself, the point of introduction of effort in the ring forming the anchor point of the reinforcement structure on this ring, and the outer end of the associated connecting rod, are all arranged in the same radial imaginary plane in which therefore is also preferably one of the structural arms connecting the housings 12 and 16. The ball joint nature of the attachment of the ends of the connecting rods 62 allows to better manage the thermal expansion of the center housing with respect to the annular structure 60 which surrounds it, and this in both the radial direction and in the longitudinal direction. Indeed, the ball joint connecting rods positioned as described above are easily able to accompany the deformations of the central casing in both directions, when it expands, without causing harmful stresses.

p0087In addition, as shown in Figure 13, each of the three rods extending in the same circumferential direction from the outer end 62b, for example clockwise as has been shown. With this configuration, in case of differential thermal expansion between the ring 60 and the casing 16 or the links 62, the ring 60 is rotatable about the central casing 16 while remaining coaxial thereto. Referring specifically to Figure 14, it is detailed the distribution of loads applied at the point of introduction efforts 68a, the principle remains similar for the other two 68b points 68c. At the point of introduction effort

p008868a, the force 76 from the reinforcing structure associated 64a is arranged substantially radially, and more particularly in the imaginary radial plane corresponding 66a. The radial force 76 is taken up firstly by a compression or tensile stresses 78 in the rod 62, and on the other hand by an effort 80 necessarily substantially tangential, in the ring 60, this effort is also said membrane effort. Therefore, at each of the three ring of force introduction points, it tends to respond to mechanical stress of the reinforcing structures by a substantially tangential force, greatly limiting the risk of ovalization.

p0089Figure 13a showing an alternative embodiment there is provided, in addition to the three links 62 described above with reference to Figure 13, a fourth rod 62 appropriately positioned.

p0090In order to homogenize the distribution of the forces, the three links 62 from the respective stress introduction points 68a, 68b, 68c is associated a fourth connecting rod 62 connecting the ring 60 to the housing 16, this fourth connecting rod being arranged symmetrically relative to that attached to the third engine, by central symmetry center constituted by the axis 5. Thus, the outer end 62b is also arranged to be traversed, in front view, for the imaginary radial plane 66c passing by the longitudinal axis 5 and the point of introduction of 68c efforts concerned.

p0091In addition, as shown in Figure 13a, each of the four rods therefore extends in the same circumferential direction from the outer end 62b, for example clockwise as has been shown. Again, in case of differential thermal expansion between the ring 60 and the casing 16 or the links 62, the ring 60 is rotatable about the central casing 16 while remaining coaxial celui- this. Moreover, in the particular case where the first and second engine mounts are arranged in the plane P2, the four links 62 tangential to the housing 16 are then distributed in a central center of symmetry formed by the longitudinal axis 5.

p0092In Figure 15 showing another embodiment, the links 62 arranged in the manner described above are always connected to the ring 60, who is not only connected to the force introduction points 68a, 68b, 68c but also by a body 86 of inner radial boundary of the annular secondary flow channel 88 (from the English "IFS" Inlet Fan structure). For information, this structure 86 is disposed radially inward relative to a 90 radial outer boundary structure of the annular secondary flow channel (from the English "OFS" Outlet Fan Structure), itself located in the rear extension of the inner skin side casings of the engine mount.

p0093In such a case, as mentioned above, one can effectively provide that reinforcing structures 64a, 64b, 64c, running along the internal structure 80, play an additional role air bifurcation in the secondary flow of the turbojet.

p0094Figures 13 to 15, the jet engine 2 is shown in a position as set when it is suspended under the wing. However, here again, the particular configuration of the mounting means 62, described above, can be considered to any positioning of the turbojet, in particular when it is reported back portion of the fuselage, as shown in Figures 8 and 9.

p0095Of course, various modifications may be made by the skilled person in the motor assembly 1 for an aircraft which has just been described, purely by way of nonlimiting example. In this regard, it is noted that the two optional specificities shown in Figures 3-9 respectively and Figures 10 to 12 have been described in combination, but could be provided one without the other, without departing from the scope 1 of the invention.