Thrust reversing device for the gondola of an aircraft turbo engine

Description

Technical field of the invention

[0001] The invention relates to the field of aircraft twin-flow turbojet nacelles, and in particular to the thrust reversers which equip these nacelles. Technical background

[0002] The technical background includes in particular documents FR-A1-2 938 878, FR-A1-2 954 410, US-A1-2020 / 003151, US-A-3,797,785 and US-B2-6,385,964.

[0003] Thrust reversers are now widely used in aircraft nacelles, particularly in nacelles housing a turbofan engine. As is known, such a turbofan engine generates, via the rotating blades of a fan, a flow of hot air (called the primary flow) from a combustion chamber, and a flow of cold air (called the secondary flow) which circulates outside the turbofan engine through an annular channel formed between a fairing of the turbofan engine and an internal wall of the nacelle. The two airflows are then ejected from the turbofan engine through the rear of the nacelle and thus generate thrust.

[0004] In such a configuration, the role of a thrust reverser is, during a landing phase of the aircraft, to improve its ground braking capacity by redirecting forward at least part of the thrust generated by the turbojet. In particular, when the thrust reverser is in action, it obstructs the annular channel of the cold air flow (i.e. the secondary flow) and directs this flow towards the front of the nacelle, thereby generating counter-thrust.

[0005] The means used to achieve this reorientation of the cold air flow vary depending on the type of reverser. However, in all cases, the structure of a thrust reverser comprises one or more movable cowls that can be moved between, on the one hand, a deployed position (also called the thrust reversal position) in which they open a passage in the nacelle intended for the diverted flow, and on the other hand, a retracted position (also called the direct jet position) in which they close this passage. The cowls can thus perform a function of activating other means of diversion such as flaps. In this case, the flaps, actuated by the movement of the movable cowls, obstruct, at least in part, the channel in which the secondary flow circulates.

[0006] Furthermore, in the case of a so-called diverter thrust reverser, the reorientation of the airflow is then carried out by diverter grilles.

[0007] A prior art inverter is illustrated in Figures 1 and 2 This inverter is of the grid inverter or cascade inverter type.

[0008] This type of reverser comprises at least one movable cowl 9 relative to a fixed structure comprising an upstream annular frame 15, the cowl 9 having an external wall 17 and an internal wall 11 intended to delimit, in a direct jet position of the turbojet ( figure 1 ), an external wall of the annular vein 6 in which the secondary flow F11 flows. The reverser further comprises flaps 10 mounted in an articulated manner on the movable cowl 9 and actuated by connecting rods 12 when the movable cowl moves downstream, so that, in a thrust reversal position ( figure 2 ), each flap 10 comprises a zone extending into the annular channel 6 so as to divert at least part of the secondary flow F11 out of the vein 6.

[0009] In the case of this type of inverter, the reorientation of the secondary flow F11 is carried out by deflection grids 13, the movable cowl 9 having only a simple sliding function aimed at uncovering or covering these grids 13, the translation of the movable cowl 9 being carried out along a longitudinal axis substantially parallel to the axis of the nacelle 1 and of the inverter.

[0010] A housing 14 is provided in the cover 9 and allows the grids 13 to be housed when the inverter is not actuated, that is to say in the direct jet position, as shown in figure 1 Alternatively, the grids 13 could be slidably mounted and housed in the fixed structure.

[0011] The grids 13 are arranged adjacent to each other in an annular area surrounding the annular channel 6, the grids 13 being arranged edge to edge so that the smallest possible gap is left between them. In this way, the vast majority of the secondary flow F11 deflected by the flaps 10 passes through the grids 13.

[0012] The grids may be fixed by their upstream ends to the structure, as illustrated in the drawings. Alternatively and as described in prior application FR-A1-3 076 865, the grids may be fixed by their downstream ends to the hood.

[0013] An annular deflection element 19, commonly called a deflection edge, covers the inner periphery of the frame 15. This deflection element 19 has a rounded cross-section and extends from the outer periphery of the annular vein 6 to the upstream end of the grids 13. The frame 15 and the deflection element 19 form a fixed annular structure.

[0014] The thrust reverser further comprises means for moving and guiding the movable cowls 9 which generally comprise actuators 21 which each have an elongated shape and extend parallel to the axis of the thrust reverser. Each actuator 21 has an upstream end 21a fixed to the fixed structure comprising the frame 15, and a downstream end 21b which is fixed to the cowl. In the aforementioned variant where the grids are slidably mounted, the actuator is mounted between the fixed structure and the grids.

[0015] THE figures 3 And 4allow you to see the shutters 10 in the first position corresponding to the figure 1 ( figure 3 ), and in the second position of diversion of the secondary flow F11 corresponding to the figure 2 ( figure 4 ).

[0016] These figures 3 And 4 allow the shape of each of these flaps 10 to be seen and also allow fairings 20 to be seen which are carried by the cover 9 and which are arranged between the flaps 10.

[0017] The fairings 20 are distributed around the axis A of the thrust reverser 30 and each have a generally triangular or truncated cone shape. Each of these fairings 20 comprises a downstream base 20a connected to the cowl, an upstream tip 20b, and two lateral edges 20c connecting the tip 20b to the base 20a.

[0018] The flaps 10 are distributed around the axis A and mounted between the fairings 20. Each flap 10 comprises: ° an upstream edge 10a intended to be located downstream of the deflection element 19 (generally with the interposition of an intermediate shroud) in the first position, ° a downstream edge 10b intended to be flush with the cowl 9 in the first position, and ° lateral edges 10c each comprising two rectilinear parts, respectively upstream 10c1 and downstream 10c2. The upstream parts 10c1 of the lateral edges 10c of the flaps 10 are intended to be flush with the upstream parts 10c1 of the lateral edges 10c of adjacent flaps 10 in the first position, and the downstream parts 10c2 of the lateral edges 10c of the flaps 10 are intended on the one hand to be flush with the lateral edges 20c of the fairings 20 in the first position ( figure 3 ), and on the other hand the downstream parts 10c2 of the lateral edges 10c of the flaps 10 in the second position ( figure 4 ).

[0019] THE figures 5 And 6are enlarged views of flaps 10 and fairings 20 and show that the flaps 10 and fairings 20 extend in line with each other in the first aforementioned position.

[0020] There Figure 5 also shows that the flaps 10 extend inside an upstream annular portion 9a of the cowl 9 and at a radial distance from this portion 9a. The cowl 9 further comprises a downstream annular portion 9b which extends in the extension of the flaps 10 and the fairings 20 towards the downstream. In other words, the downstream bases 20a of the fairings 20 are connected to an upstream end 9ba of this downstream portion 9b of the cowl 9, and the downstream edges 10b of the flaps 10 are flush with this upstream end 9ba of the downstream portion 9b of the cowl 9 in the first position mentioned above.

[0021] The flaps 10 define with the upstream portion of the hood 9 an annular space E.

[0022] In operation, air from the secondary flow F11 can enter this space E through the gap E1 formed between the upstream edges 10a of the flaps 10 and the deflection element 19 (or between the upstream edges 10a and the intermediate shell interposed between the flaps 10 and the deflection element 19). This air circulates in the space E, then is ejected through the gaps E2, E3 formed on the one hand between the lateral edges 10c of the flaps 10 and the lateral edges 10c, 20c of the adjacent flaps 10 ( figure 8 ) or fairings 20 ( figure 7 ), and between the downstream edges 10b of the flaps 10 and the upstream end 9ba of the downstream portion 9b of the cowl 9. When this air is ejected, vortex structures form in the secondary vein 6, generating thrust losses, which is problematic.

[0023] A first solution to this problem would be to reduce the dimensions of the gaps E1, E2 and E3 as much as possible. However, managing these dimensions is becoming increasingly complex through the use of composite materials for the flaps. These materials are relatively flexible and tend to deform during operation. It is therefore very difficult to control the dimensions of the gaps, particularly for large-diameter turbojets.

[0024] A second solution would consist of providing seals upstream of the flaps 10, between the upstream edges 10a of these flaps and the deflection element 19, to prevent air from entering the space E. However, this solution would also not be satisfactory because the seals would increase the forces applied by the actuator 21 on the deflection element 19. The seals would have to be of relatively large dimensions to compensate for the relative movements between the fixed structure and the moving elements, which would increase their mass. Furthermore, in the thrust reversal position, the presence of the seals upstream of the flaps 10 could hinder the opening kinematics of the thrust reverser by coming into contact with the cowl 9 or other elements of the turbojet engine in the second thrust reversal position.

[0025] A third solution would consist of providing flaps 10 with walls equipped with elastically deformable lateral extensions. These flexible lateral extensions would extend in the extension of the main wall of each flap and would define the lateral edges of this flap. These extensions would be intended to cooperate in sealing with the elements located in the vicinity of the flap 10. However, this solution would not be satisfactory either. In the thrust reversal position, these flexible extensions would be subjected to the unsteady reverse flow, which would damage them. These extensions would have to be integrated into the flaps 10 and would not be easily replaceable. Furthermore, the presence of these lateral extensions would result in a reduction of the acoustic surfaces of the flaps 10 and in risks of contact and damage with the cowl 9 or other elements of the turbojet in the second thrust reversal position.

[0026] The invention provides a simple, effective and economical solution to the problem mentioned above. Summary of the invention

[0027] The present invention provides a thrust reverser for a nacelle of an aircraft twin-flow turbojet engine, this thrust reverser having a generally annular shape around an axis and comprising: a fixed upstream annular structure which extends around an annular flow vein of a secondary flow, deflection grilles which are distributed around the axis, at least one annular cowl movable in axial translation relative to the fixed structure, from an upstream position in which the grilles are covered to a downstream thrust reversal position in which the grilles are uncovered, this cowl comprising an upstream annular portion and a downstream annular portion of the fairings carried by the cowl, these fairings being distributed around the axis, each of these fairings extending inside and at a distance from the upstream portion of the cowl and comprising a downstream base connected to an upstream end of the downstream annular portion of the cowl, an upstream tip, and two lateral edges connecting the tip to the base, flaps for deflecting the secondary flow through the grilles, the flaps being distributed around the axis and mounted between the fairings,the flaps being movable from a first position in which they extend inside and at a distance from the upstream portion of the cowl and in the extension of the adjacent fairings, when the cowl is in the upstream position, and a second position in which they extend into the vein to deflect the secondary flow, when the cowl is in its downstream position, each of these flaps comprising: ° an upstream edge intended to be flush with said upstream structure in the first position, ° a downstream edge intended to be flush with the upstream end of the downstream portion of the cowl in the first position, and ° lateral edges each comprising two parts, respectively upstream and downstream, the upstream parts of the lateral edges of the flaps being intended to be flush with the upstream parts of the lateral edges of adjacent flaps in the first position, and the downstream parts of the lateral edges of the flaps being intended on the one hand to be flush with the lateral edges of the fairings in the first position,and on the other hand the downstream parts of the lateral edges of the flaps in the second position, and a device for moving the hood and the flaps, or even the grilles, , characterized in that each of the flaps comprises a rigid wall extending to the upstream, downstream and lateral edges, and defining these edges, this wall comprising an aerodynamic internal face and an external face facing the upstream portion of the cowl in the first position, and in that the reverser further comprises: first sealing elements located on the external face of the wall of each flap and / or on an external face of the fairings, these first sealing elements extending along the lateral edges of the flap and / or of the fairings so as to ensure a seal along these edges in the first position.

[0028] The invention makes it possible to limit the air flows from the aforementioned space, located between the flaps and the upstream portion of the cowl, through the lateral gaps located between the lateral edges of the flaps and the lateral edges of the fairings and adjacent flaps. The sealing elements allow, for example, the air contained in this space to lose its energy, or to be diverted or blocked, before reaching the gaps.

[0029] In this application, the term "rigid" means an element or material that is not flexible or elastically deformable. A rigid element or material is therefore an element or material that is not elastically deformable. Conversely, the term "flexible" means an element or material that is elastically deformable.

[0030] In the present application, the term "sealing" means reducing or even eliminating gas leaks between two elements. In the event of sealed contact between these two elements, gas leaks are prevented. In the absence of contact, for example in the case where the clearance between the elements is reduced to reduce leaks between these elements, the sealing would be provided by the energy loss of the gas flow between the elements.

[0031] The inverter according to the invention may comprise one or more of the following features, taken in isolation from one another or in combination with one another: --- the fairings each have a generally triangular or truncated cone shape; -- in the aforementioned upstream position, the grids are covered by the upstream annular structure or the annular cover; -- each of these fairings extends in the continuity of the downstream annular portion of the cover, and is therefore integrated into this downstream portion; -- the displacement device comprises at least one actuator; -- the device or the actuator comprises a first axial end, in particular upstream, fixed to said structure, and a second axial end, in particular downstream, fixed to said cover; -- the grids comprise upstream ends fixed to said structure or downstream ends fixed to said cover; each of the flaps is devoid of a sealing element along its upstream edge; the first sealing elements are added and fixed on the external face of the wall of each flap and / or on the external face of the fairings;the first sealing elements are integrated into the wall of the flap and / or the fairing; -- the first sealing elements are at least partly (or entirely) rigid; the first sealing elements comprise: a first elastically deformable longitudinal member which extends along the lateral edge of the flap, and a second rigid longitudinal member which extends along the lateral edge of the fairing and which cooperates by sealing support with said first member, in said first position, or conversely, a first elastically deformable longitudinal member which extends along the lateral edge of the fairing, and a second rigid longitudinal member which extends along the lateral edge of the flap and which cooperates by sealing support with said first member, in said first position; -- the first member has a circular or elliptical cross-section; -- the first member comprises at least one sealing lip;said sealed support is a radially oriented support or a circumferentially oriented support relative to said axis; the second member has an L-shaped or parallelepiped shape in cross-section; the second member comprises a longitudinal profile extending along the edge on which it is fixed and generating a radial excess thickness on this edge; the profile comprises a longitudinal rim forming a lateral extension of the edge on which the second member is fixed;the first sealing elements comprise: a first rigid longitudinal member which extends along the lateral edge of the flap, and optionally a second rigid longitudinal member which extends along the lateral edge of the fairing and which is separated by a predetermined clearance from said first member, in said first position, or conversely, a first rigid longitudinal member which extends along the lateral edge of the fairing, and optionally a second rigid longitudinal member which extends along the lateral edge of the flap and which is separated by a predetermined clearance from said first member, in said first position; said clearance is a radial clearance or a circumferential clearance relative to said axis;each of the flaps comprises: at least one second sealing element located on the external face of the wall and / or on the upstream end of the downstream portion of the cover, this second sealing element extending along the downstream edge of the flap so as to ensure sealing or to reduce leaks along this edge in the first position; the second element is identical to said first elements; the second sealing element is added and fixed on the external face and / or on the upstream end of the downstream portion of the cover; the second sealing element is integrated into the wall of the flap and / or on the upstream end of the downstream portion of the cover; -- the second sealing element is at least partly (or entirely) rigid. ;

[0032] The present invention also relates to an aircraft propulsion assembly comprising a dual-flow turbojet engine and a thrust reverser as described above. Brief description of the figures

[0033] The present invention will be better understood and other details, characteristics and advantages of the present invention will appear more clearly on reading the description of a non-limiting example which follows, with reference to the appended drawings in which: [ Fig. 1 ] there figure 1 is a partial schematic view, in longitudinal section, of a thrust reverser according to the prior art in the direct jet position; [ Fig.2 ] there figure 2 is a partial schematic view, in longitudinal section, of the thrust reverser of the figure 1 in reverse thrust position; [ Fig.3 ] there figure 3 is a schematic perspective view of a thrust reverser, in the direct jet position; [ Fig.4 ] there figure 4 is a schematic perspective view of the thrust reverser of the figure 3 , in the reverse thrust position; [ Fig.5 ] there Figure 5is a partial schematic and perspective view of a thrust reverser and shows flaps and fairings of this reverser, in the direct jet position; [ Fig.6 ] there figure 6 is a schematic sectional view along line VI-VI of the Figure 5 ; [ Fig.7 ] there figure 7 is a schematic sectional view along line VII-VII of the Figure 5 ; [ Fig.8 ] there figure 8 is a view similar to that of the Figure 5 and represents a first embodiment of the invention; [ Fig.9 ] there figure 9 is a view similar to that of the Figure 5 and represents a second embodiment of the invention; [ Fig. 10 ] there figure 10 is a view similar to that of the Figure 5 and represents a third embodiment of the invention; and [ Fig. 11 ] there figure 11 is a view similar to that of the Figure 5 and represents a fourth embodiment of the invention. Detailed description of the invention

[0034] THE figures 1 to 7 have been described in the above.

[0035] We now refer to Figures 8 to 12 which illustrate several embodiments of a thrust reverser 30 according to the invention for an aircraft turbojet nacelle.

[0036] The foregoing description in relation to the figures 1 to 7 applies to the thrust reverser 30 according to the invention.

[0037] This inverter 30 includes in particular: - a fixed upstream annular structure which extends around an annular flow vein of a secondary flow, this structure comprising the casing 15 and the deflection element 19 in the example shown in Figures 1 and 2 , - deflection grids 13 which are distributed around the axis A of the inverter 30, - at least one annular cover 9 movable in axial translation relative to the fixed structure, from an upstream position ( figures 1 , 3 , 5 And 6) in which the grids 13 are covered, up to a downstream thrust reversal position ( figures 2 And 4 ) in which the grids 13 are uncovered, this cover 9 comprising an upstream annular portion 9a and a downstream annular portion 9b, - fairings 20 carried by the cover 9, these fairings 9 being distributed around the axis A and each having a generally triangular or truncated cone shape in the example shown, each of these fairings 20 extending inside and at a distance from the upstream portion 9a of the cover 9 and comprising a downstream base 20a connected to an upstream end 9ba of the downstream annular portion 9b of the cover 9, an upstream point 20b, and two lateral edges 20c connecting the point 20b to the base 20a, - flaps 10 for deflecting the secondary flow F11 through the grids 13, the flaps 10 being distributed around the axis A and mounted between the fairings 20, the flaps 10 being movable from a first position ( figures 1 , 3 , 5And 6 ) in which they extend inside and at a distance from the upstream portion 9a of the cowl and in the extension of the adjacent fairings 20, when the cowl 9 is in the upstream position, and a second position ( figures 2 And 4 ) in which they extend into the vein 6 to divert the secondary flow F11, when the cover 9 is in its downstream position.

[0038] Each of the 10 sections includes: ° an upstream edge 10a intended to be flush with the upstream structure in the first position, ° a downstream edge 10b intended to be flush with the upstream end 9ba of the downstream portion 9b of the cowl 9 in the first position, and ° lateral edges 10c each comprising two parts, respectively upstream 10c1 and downstream 10c2. These upstream parts 10c1 of the lateral edges 10c of the flaps 10 are intended to be flush with the upstream parts 10c1 of the lateral edges 10c of adjacent flaps 10 in the first position, and the downstream parts 10c2 of the lateral edges 10c of the flaps 10 being intended on the one hand to be flush with the lateral edges 20c of the fairings 20 in the first position, and on the other hand the downstream parts 10c2 of the lateral edges 10c of the flaps 10 in the second position.

[0039] The inverter 30 further comprises actuators 20 for moving the hood 9 and the flaps 10 between the aforementioned positions.

[0040] Each of the flaps 10 comprises a rigid wall 32 extending to the upstream 10a, downstream 10b and lateral 10c edges, and defining these edges. This wall 32 comprises an aerodynamic internal face 32a and an external face 32b facing the upstream portion 9a of the cover 9 in the first position. This wall 32 is for example made of aluminum, or thermosetting composite, or thermoplastic composite.

[0041] According to the invention, the inverter 30 further comprises sealing elements located on the external face 32b of the wall 32 of each flap 10 and / or on an external face 20d of the fairings 20.

[0042] The sealing elements preferably extend along peripheral edges of the flaps 10 and the fairings 20. The sealing elements extending along the lateral edges 10c are called first sealing elements, referenced 34, and the sealing elements extending along the lateral edges 20c of the fairings 20 are called first sealing elements, referenced 36. These first sealing elements 34, 36 ensure sealing by loss of energy of the lateral flows along these lateral edges 10c, 20c in the first position.

[0043] The second sealing elements, referenced 38, are the sealing elements extending along the downstream edges 10b of the flaps, and the second sealing elements, referenced 40, are the sealing elements extending along the edges opposite the upstream end 9ba of the downstream portion 9b of the cover 9. These second sealing elements 38, 40 ensure sealing along these edges 10b in the first position.

[0044] THE figures 8 to 11illustrate several embodiments of these sealing elements 34 to 40. Although the figures show the particular case of the seal (and the sealing elements 34, 36) between the edge 10c of a flap 10 and the edge 20c of a fairing, these figures and the associated description are transposable to the case of the seal (and the sealing elements 34, 36) between two edges 10c of two adjacent flaps 10, and of the seal (and the sealing elements 38, 40) between the edge 10b of a flap and the upstream end 9ba of the downstream portion 9b of the cowl 9.

[0045] In the first embodiment shown in the figure 8 , the first sealing elements 34 of the flap 10 comprise a first elastically deformable longitudinal member 42 which extends along the lateral edge 10c of the flap 10.

[0046] The first sealing elements 36 of the fairing 20 comprise a second rigid longitudinal member 44 which extends along the lateral edge 20c of the fairing 20 and which cooperates by sealing support with the member 42, in said first position.

[0047] The reverse configuration of the figure 8 is however conceivable. The member 42 could extend along the lateral edge 20c of the fairing 20, and the member 44 could extend along the lateral edge 10c of the flap 10.

[0048] Although this is not limiting, the first member 42 has a circular or elliptical cross-section, in particular when at rest in the unstressed state. This member 42 preferably has a solid section and is for example made of elastomer. In the example shown, the sealed support C1 between the members 42, 44 is a support oriented radially relative to the axis A of the inverter 30.

[0049] The second member 44 comprises a longitudinal profile extending along the edge 20c and generating a radial excess thickness on this edge. This member 44 or this profile comprises a longitudinal rim 46 which forms a lateral extension of the edge 20c and which externally covers the edge 10c of the flap in order to bear in the radial direction on the member 42. In this embodiment, the member 44 or the profile has a general L-shape in cross section.

[0050] The method of realization of the figure 9 differs from the previous embodiment in that the member 44 does not include a rim 46 and the sealed support C2 between the members 42, 44 is a support oriented circumferentially relative to the axis A of the inverter 30. A longitudinal side of the member 44 or of the profile bears in the circumferential direction on the member 42. In this embodiment, the member 44 or the profile has a generally parallelepiped shape in cross section.

[0051] The reverse configuration of the figure 9 is also conceivable. The member 42 could extend along the lateral edge 20c of the fairing 20, and the member 44 could extend along the lateral edge 10c of the flap 10.

[0052] The method of realization of the figure 10 differs from the first embodiment in that there is no member 42. Only the member 44 is present on the fairing 20. There is also no sealed support between the first sealing elements.

[0053] The seal between the fairing and the flap is here ensured by a predetermined clearance between the member 44 and the lateral edge 10c of the flap. In the example shown, this clearance R1 is oriented in the radial direction.

[0054] The member 44 also has an L-shaped section and includes a rim 46 which covers the lateral edge 10c of the flap and which is separated from this edge by the radial clearance R1. The reverse configuration of the figure 10is also conceivable. The member 44 could extend along the lateral edge 10c of the flap 10 and define a radial clearance R1 with the edge 20c of the fairing.

[0055] In the embodiment of the figure 11 , the first sealing elements 34 of the flap 10 comprise a first rigid longitudinal member 48 which extends along the lateral edge 10c of the flap 10.

[0056] The first sealing elements 36 of the fairing 20 comprise a second rigid longitudinal member 44 which extends along the lateral edge 20c of the fairing 20.

[0057] In this example, each of the members 42, 44 comprises a longitudinal profile extending along the corresponding edge and generating a radial excess thickness on this edge. Each member 42, 44 or profile has a generally parallelepiped shape in cross section.

[0058] The seal between the fairing 20 and the flap 10 is here ensured by a predetermined clearance between the members 42, 44. In the example shown, this clearance R2 is oriented in the circumferential direction.

[0059] In the different embodiments of the figures 8 to 11, the rigid materials used for the sealing elements can be a metal, a composite or a thermoplastic, Teflon ®< , etc. The flexible materials used can be an elastomer or rubber. In the presence of a flexible sealing element, the seal is obtained by support, this support being able to be maintained when the parts undergo relative movements. In the absence of a flexible sealing element, the seal is obtained by the creation of a baffle or obstacle limiting the passage of air. The optimization of the dimensions (L, R1, R2, H1, H2...), of the organs and profiles, and in particular of their parts defining baffles or overlaps, makes it possible to optimize the seal of the assembly. Furthermore, the absence of contact between the parts makes it possible to limit the risk of wear in particular.

[0060] The invention is particularly advantageous because it makes it possible to solve the problem mentioned above, without affecting the aerodynamic internal faces of the flaps which must be defined by rigid walls so as not to alter the deflection and acoustic performance of the flaps.

Claims

1. A thrust reverser (30) for a nacelle of an aircraft turbofan engine, this thrust reverser having a generally annular shape about an axis (A) and comprising: - a stationary upstream annular structure (15, 19) which extends around an annular duct (6) for a secondary flow (F11), - deflection grids (13) distributed around the axis, - at least one annular cover (9) movable in axial translation relative to the stationary structure, from an upstream position in which the grids (13) are covered, to a downstream thrust-reversal position in which the grids are uncovered, this cover (9) comprising an upstream annular segment (9a) and a downstream annular segment (9b), - fairings (20) carried by the cover (9), these fairings being distributed around the axis (A), each of these fairings extending inside and at a distance from the upstream segment (9a) of the cover (9) and comprising a downstream base (20a) connected to an upstream end (9ba) of the downstream annular segment (9b) of the cover (9), an upstream tip (20b), and two lateral edges (20c) connecting the tip to the base, - deflection flaps (10) for deflecting the secondary flow (F11) through the grids (13), the flaps being distributed around the axis (A) and mounted between the fairings (20), the flaps being movable from a first position in which they extend inside and at a distance from the upstream segment (9a) of the cover (9) and in the extension of the adjacent fairings (20), when the cover (9) is in the upstream position, and a second position in which they extend into the duct (6) to deflect the secondary flow (F11), when the cover (9) is in its downstream position, each of these flaps (10) comprising: • an upstream edge (10a) designed to be flush with said upstream structure in the first position, • a downstream edge (10b) intended to be flush with the upstream end (9ba) of the downstream segment (9b) of the cover (9) in the first position, and • lateral edges (10c) each comprising two portions, respectively upstream (10c1) and downstream (10c2), the upstream portions (10c1) of the lateral edges (10c) of the flaps (10) being intended to be flush with the upstream portions (10c1) of the lateral edges of adjacent flaps in the first position, and the downstream portions (10c2) of the lateral edges (10c) of the flaps (10) being intended on the one hand to be flush with the lateral edges (20c) of the fairings (20) in the first position, and on the other hand the downstream portions (10c2) of the lateral edges (10c) of the flaps (10) in the second position, - and a moving device (21) for moving the cover (9) and the flaps (10), wherein each of the flaps (10) comprises a rigid wall (32) extending as far as the upstream (10a), downstream (10b) and lateral (10c) edges, and defining these edges, this wall (32) comprising an aerodynamic internal face (32a) and an external face (32b) facing the upstream segment (9a) of the cover (9) in the first position, and in that the reverser (30) further comprises: - first sealing elements (34, 36) located on the external face (32b) of the wall (32) of each flap (10) and / or on an external face (20d) of the fairings (20), these first sealing elements extending along the lateral edges (10c, 20c) of the flap (10) and / or of the fairings (20) so as to provide a sealing along these edges in the first position.

2. The thrust reverser (30) according to claim 1, wherein the first sealing elements (34, 36) comprise: - a first elastically deformable longitudinal member (42) which extends along the lateral edge (10c) of the flap (10), and a second rigid longitudinal member (44) which extends along the lateral edge (20c) of the fairing (20) and which cooperates by sealed support with said first member (42), in said first position, - or conversely, a first elastically deformable longitudinal member which extends along the lateral edge of the fairing, and a second rigid longitudinal member which extends along the lateral edge of the flap and which cooperates by sealed support with said first member, in said first position.

3. The thrust reverser (30) according to claim 2, wherein said sealed support is a radially oriented support or a circumferentially oriented support with respect to said axis.

4. The thrust reverser (30) according to claim 2 or 3, wherein the second member (44) is L-shaped or parallelepipedal in cross-section.

5. The thrust reverser (30) according to one of claims 2 to 4, wherein the second member (44) comprises a longitudinal profile extending along the edge to which it is attached and generating a radial excessive thickness on this edge.

6. The thrust reverser (30) according to claim 5, wherein the profile comprises a longitudinal rim (46) forming a lateral extension of the edge to which the second member (44) is attached.

7. The thrust reverser (30) according to claim 1, wherein the first sealing elements comprise: - a first rigid longitudinal member (48) which extends along the lateral edge (10c) of the flap (10), and optionally a second rigid longitudinal member (44) which extends along the lateral edge (20c) of the fairing (20) and which is separated by a predetermined clearance (R1, R2) from said first member (48), in said first position, - or conversely, a first rigid longitudinal member which extends along the lateral edge of the fairing, and optionally a second rigid longitudinal member which extends along the lateral edge of the flap and which is separated by a predetermined clearance from said first member, in said first position.

8. The thrust reverser (30) according to claim 7, wherein said clearance is a radial clearance (R1) or a circumferential clearance (R2) with respect to said axis (A).

9. The thrust reverser (30) according to one of the preceding claims, wherein each of the flaps (10) comprises: - a second sealing element (38, 40) located on the external face (32b) of the wall (32) and / or on the upstream end (9ba) of the downstream segment (9b) of the cover (9), this second sealing element extending along the downstream edge (10b) of the flap so as to seal or reduce the leakages along this edge in the first position.

10. The thrust reverser (30) according to claim 9, wherein the second sealing element is fitted and attached to the external face and / or to the upstream end of the downstream segment of the cover.

11. The thrust reverser (30) according to claim 9, wherein the second sealing element is integrated into the wall of the flap and / or on the upstream end of the downstream segment of the cover.

12. The thrust reverser (30) according to one of the preceding claims, wherein each of the flaps has no sealing element along its upstream edge.

13. The thrust reverser (30) according to one of the preceding claims, wherein the first sealing elements are fitted and attached to the external face of the wall of each flap and / or to the external face of the fairings.

14. The thrust reverser (30) according to one of claims 1 to 12, wherein the first sealing elements are integrated into the wall of the flap and / or into the fairing.

15. A propulsion unit for an aircraft, comprising a turbofan engine and a thrust reverser (30) according to one of the preceding claims.