Cascade for thrust reverser comprising blade stages connected by means of honeycomb structure thereof

Abstract

A blade cascade (22) for a cascade thrust reverser (19) of an aircraft propulsion assembly nacelle, comprising: a first blade stage (22A) and a second blade stage (22B) stacked, each blade stage comprising a plurality of rows of blades (50A, 50B) and spars (52A, 52B), the blades and spars intersecting each other thereby forming a honeycomb structure (53A, 53B) for passing air flow; and a leading end flange (40) and a trailing end flange (42) each rigidly connected to all or some spars of at least one of the blade stages. The honeycomb structures are connected to each other independently of the front end flange and the rear end flange by connecting means (66) configured to prevent relative displacement between the honeycomb structures. This results in good thrust reversion performance, compatibility with inexpensive manufacturing processes, and reduced vibration instability and aerodynamic disturbances.

Description

Technical Field

[0001] The present invention relates to the field of aircraft propulsion assembly nacelles and, more particularly, to cascade vanes for a thrust reverser and to a thrust reverser comprising such a cascade vane. Background Art

[0002] A thrust reverser is a device used to deflect the air flow (usually the secondary flow) forward through the propulsion assembly in order to shorten landing distance and limit brake loads on the landing gear.

[0003] A cascade reverser generally comprises a cascade of vanes for deflecting the air flow and integrated into a fixed structure of the reverser for connection to the turbine engine casing. The movable structure of the reverser comprises one or more reverser covers and is mounted so as to be translatable relative to the fixed structure between an extended position, referred to as a "direct jet" configuration, and a retracted position, referred to as a "reverse jet" configuration. In the direct jet configuration, the cascade is arranged in a housing defined by the reverser cover, so that the cascade is isolated from the secondary flow path of the propulsion assembly by the radial inner wall of the reverser cover. On the other hand, in the reverse jet configuration, the reverser cover is axially offset relative to the cascade, so that the cascade is exposed to the air flow to be deflected on the inside and to the external environment on the opposite side, so that the air flow can be redirected into the external environment.

[0004] A thrust reverser cascade typically includes a front flange and a rear flange for securing the cascade to a fixed or movable structure of the nacelle. Such a cascade also includes a plurality of rows of blades arranged between the two flanges and shaped to deflect the air flow toward the front of the cascade, and a spar connecting the flanges to each other. The rows of blades intersect the spar, thereby together forming a honeycomb structure that defines a plurality of cells that form as many channels as possible for the air flow. Typically, the blades extend perpendicular to the spar. The spar typically extends in a direction in an axial plane, while the blades typically extend in a circumferential direction relative to the axis of the propulsion assembly. However, other orientations are also possible.

[0005] The spar absorbs most of the mechanical loads of the honeycomb structure, while the main function of the blades is to deflect the air flow to achieve thrust reversal.

[0006] Therefore, the ability of such a cascade to effectively deflect airflow depends particularly on the curvature of the blades. Generally speaking, the more pronounced the blade curvature, the higher the cascade's aerodynamic performance. However, significant curvature makes blade manufacturing difficult, or in some cases, impossible. For example, existing cascades are produced by injection molding; while this manufacturing method offers the advantage of low cost, it tends to limit the achievable curvature of the blades due to mold release issues.

[0007] To solve these problems, cascades with two blade stages have been proposed. However, in operation, these cascades cause problems with vibration instabilities and aerodynamic disturbances. Summary of the Invention

[0008] The subject of the invention is a cascade which offers good aerodynamic performance, is at the same time easy and inexpensive to produce and is stable in operation with respect to vibrations and aerodynamic disturbances.

[0009] To this end, the present invention relates to a cascade of thrust reversers for a cascade of a nacelle of an aircraft propulsion assembly, the cascade comprising:

[0010] at least a first and a second blade stage, the first and second blade stages being superposed and each comprising a plurality of rows of blades shaped so as to deflect the air flow, and spars intersecting the blades within each blade stage so as to form, together with the blades, respective honeycomb structures for the passage of the air flow;

[0011] a front flange and a rear flange, which are arranged at the front and rear ends of the cascade, respectively, and are each rigidly connected to all or some of the spars of at least one of the blade stages;

[0012] According to the present invention, the respective honeycomb structures of the blade stage are connected to each other independently of the front flange and the rear flange by connecting means which are configured to prevent relative displacement between the honeycomb structures.

[0013] The invention makes it possible to improve the performance of thrust reverser cascades while ensuring compatibility with inexpensive manufacturing processes and, thanks to the connection between the honeycomb structures, avoids or limits the problems of vibration instability and aerodynamic disturbances encountered with two-stage cascades according to the prior art. Furthermore, this connection device makes it possible to superimpose stages of blades of different lengths by not directly connecting one or more stages to each of the front and rear flanges, which opens up many options for optimizing the reverser cascades.

[0014] In some embodiments of the invention, the connecting means comprise a first extension of the spar of the first blade stage and a second extension of the spar of the second blade stage, each of the first extensions being arranged facing a respective second extension such that the first extension and the second extension form a mutual stop which counteracts a relative displacement between the honeycomb structures.

[0015] Preferably, the connection means comprises inter-spar engagement means formed on the first extension and the second extension and co-operating to prevent relative displacement between the honeycomb structures.

[0016] In some embodiments of the invention, the connecting means comprises mutual blade engagement means formed on the blades of the first blade stage and on the blades of the second blade stage and cooperating to prevent relative displacement between the honeycomb structures.

[0017] Preferably, the mutual blade engagement means comprises a tongue fixed to a blade of one of the blade stages and extending in the direction of the other blade stage, and a groove formed in the blade of the other blade stage in which the tongue is received.

[0018] In some embodiments of the invention, the connecting means comprise means for fixing the respective honeycomb structures of the blade stages.

[0019] In some embodiments of the present invention, the first blade stage is an inner blade stage, the second blade stage is an outer blade stage, the outer blade stage is arranged on the inner blade stage, and at least one of the front end and the rear end of the outer blade stage is offset from the corresponding front end or rear end of the inner blade stage in a direction from the front flange to the rear flange.

[0020] In some embodiments of the invention, the blades of the second blade stage extend in line with corresponding blades of the first blade stage.

[0021] In other embodiments of the invention, the blades of the second blade stage are offset relative to corresponding blades of the first blade stage according to a direction from the leading edge to the trailing flange.

[0022] In some embodiments of the invention, the blades of the first blade stage comprise respective trailing edges extending beyond respective leading edges of the blades of the second blade stage in a direction from the first blade stage to the second blade stage.

[0023] In some embodiments of the invention, the spacing between consecutive rows of blades of the first blade stage is different from the spacing between consecutive rows of blades of the second blade stage.

[0024] The invention also relates to a cascade thrust reverser for a nacelle of an aircraft propulsion assembly, comprising at least one cascade vane of the type described above. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] The invention will be better understood and further details, advantages and features of the invention will become apparent on reading the following description given by way of non-limiting example and with reference to the accompanying drawings, in which:

[0026] FIG1 is a schematic perspective view of an aircraft propulsion assembly including a cascade thrust reverser according to the prior art, shown in a direct jet configuration;

[0027] FIG2 is a view similar to FIG1 showing the thrust reverser in a reverse jet configuration;

[0028] FIG3 is a schematic axial cross-sectional view of a thrust reverser of the propulsion assembly of FIG1 shown in a direct jet configuration;

[0029] FIG4 is a view similar to FIG3 showing the thrust reverser in a reverse jet configuration;

[0030] FIG5 is a schematic diagram of a blade cascade of a thrust reverser of the propulsion assembly of FIG1 ;

[0031] [ Figure 6 ] is a schematic perspective view of a cascade blade for a thrust reverser according to a preferred embodiment of the present invention;

[0032] [ Figure 7 ] is a schematic cross-sectional view of a cascade for a thrust reverser according to a preferred embodiment of the present invention;

[0033] [ Figure 8 ] is a schematic cross-sectional view of a cascade for a thrust reverser according to another preferred embodiment of the present invention;

[0034] [ Figure 9 ] is a schematic axial cross-sectional view of a blade cascade for a thrust reverser according to another preferred embodiment of the present invention;

[0035] [ Figure 10 ] is a schematic axial cross-sectional view of a blade cascade for a thrust reverser according to another preferred embodiment of the present invention;

[0036] [ Figure 11A ] is a schematic axial cross-sectional view of a blade pair according to a preferred embodiment of the present invention, the blade pair being respectively intended to be part of two superposed blade stages of a thrust reverser cascade and shown in exploded form;

[0037] [ Figure 11B ] is similar to Figure 11A , showing a pair of assembled blades;

[0038] [ Figure 11C ] is similar to Figure 11B , showing a pair of blades according to another preferred embodiment of the invention, each intended to be part of two superposed blade stages of a thrust reverser cascade;

[0039] [ Figure 12 ] is similar to Figure 9 A view showing a cascade for a thrust reverser, wherein the blades are formed similar to Figure 11C The leaf pairs;

[0040] [ Figure 13 ] is similar to Figure 12 , showing a cascade blade for a thrust reverser according to another preferred embodiment of the present invention;

[0041] [ Figure 14A ] is a very schematic axial cross-sectional view of a cascade for a thrust reverser according to a preferred embodiment of the present invention;

[0042] [ Figure 14B ]yes Figure 14A A very schematic top view, i.e. a radially external view, of a cascade of blades for a thrust reverser;

[0043] [ Figure 15 ] is a schematic perspective view of a blade pair according to another preferred embodiment of the present invention, the blades of the pair being respectively intended to be part of two superposed blade stages of a thrust reverser cascade;

[0044] [ Figure 16A ] is a schematic axial cross-sectional view of a thrust reverser for a propulsion assembly, the thrust reverser including cascade blades according to a preferred embodiment of the present invention;

[0045] [ Figure 16B ] is a schematic axial cross-sectional view of a thrust reverser for a propulsion assembly, the thrust reverser including a cascade blade according to another preferred embodiment of the present invention;

[0046] [ Figure 17 ] is a schematic perspective view of a blade cascade for a thrust reverser according to another preferred embodiment of the present invention.

[0047] Throughout the drawings, the same reference numbers may identify the same or similar elements. DETAILED DESCRIPTION

[0048] 1 and 2 show a propulsion assembly 10, which generally comprises a turbojet engine (not shown in the figures), for example of the turbofan type, surrounded by a nacelle 12. The nacelle comprises, in a manner known per se, an air inlet 14, an intermediate section 16 and an aft section 18 incorporating a thrust reverser 19 of the cascade type. FIG1 shows the nacelle 12 in a direct jet configuration, i.e. with the thrust reverser retracted, while FIG2 shows the nacelle in a reverse jet or thrust reverse configuration, i.e. with the thrust reverser deployed. Thus, FIG2 shows the cowl 20 of the aft section 18 in a retracted position and a set of thrust reverser cascades 22, distributed around the axis A of the nacelle and, more generally, around the axis A of the propulsion assembly.

[0049] Throughout this description, the axial direction X is the direction of the axis A, the vertical direction Z is the direction perpendicular to the axial direction X and is arranged so as to be oriented along a vertical line when the propulsion assembly 10 is equipped with an aircraft parked on the ground, and the transverse direction Y is perpendicular to the two preceding directions. Furthermore, the radial direction R and the circumferential direction C, or azimuthal directions, are defined with reference to the axis A, the radial direction R being a direction perpendicular to and passing through the axis A at all points, and the circumferential direction C being perpendicular to both the radial direction R and the axis A at all points. Finally, the "upstream" and "forward" directions, on the one hand, and the "downstream" and "rearward" directions, on the other hand, are defined relative to the general direction of the airflow in the turbojet engine, from upstream or forward to downstream or rearward, along the axis A.

[0050] Figures 3 and 4 show such a thrust reverser 19 in more detail in a direct jet configuration and a reverse jet configuration, respectively.

[0051] The thrust reverser 19 is usually arranged downstream of the fan casing 23 of the turbojet engine and the associated fan cowl 24 (Figures 1 and 2), which is part of the intermediate section 16 of the nacelle. The reverser 19 comprises at least one cowl, for example a cowl 20, which is axially movable between a forward or retracted position corresponding to the direct jet configuration and a rearward or extended position corresponding to the reverse jet configuration. This cowl 20 has an outer wall 26 and an inner wall 28, which, in the direct jet configuration (Figure 3), are each intended to be inscribed in the aerodynamic shell of the nacelle and to externally delimit an annular channel 30 in which the secondary flow SF of the turbojet engine flows. The reverser 19 comprises at least one reverse flap 32, which is hingedly mounted on the inner wall 28 of the cowl 20 and is actuated by at least one connecting rod 34 when the cowl 20 moves rearward, so that, in the reverse jet configuration ( FIG. 4 ), at least a portion of each reverse flap 32 extends in the annular channel 30 in order to deflect at least a portion of the secondary flow SF from the annular channel 30 towards the cascade vanes 22. The connecting rod 34 is, for example, hinged on a fixed internal structure 36 of the nacelle which internally delimits the annular channel 30.

[0052] Each movable cowl 20 comprises at least one housing 38 defined between the outer wall 26 and the inner wall 28 of the cowl and making it possible to house one or more of the cascade vanes 22 in the direct injection configuration ( FIG. 3 ).

[0053] One of the blade rows 22 can be seen in Figures 3 and 4 and is shown here in a known configuration. Each of these blade rows 22 has a substantially perforated plate shape that is curved in the circumferential direction C.

[0054] Each cascade blade 22 comprises in particular a front flange 40 and a rear flange 42 attached to a front frame 44 and a rear frame 46 , respectively, each secured to a fixed structure of the nacelle, such as the fan case 23 .

[0055] Furthermore, each cascade includes a plurality of rows of blades 50 ( FIGS. 2-4 ) arranged between the two flanges 40 , 42 and shaped to deflect the airflow forward, and a spar 52 ( FIG. 2 ). The blades 50 are arranged between the two flanges 40 , 42 and are shaped to deflect the airflow forward. Each spar 52 rigidly connects the leading flange 40 to the trailing flange 42 ( FIGS. 3-4 ). The rows of blades 50 intersect the spars 52 so that they together form a honeycomb structure 53 that defines a plurality of grooves 54 , each of which forms a channel for the airflow. The spars 52 extend in a direction from the leading flange 40 to the trailing flange 42 , which is generally parallel to the axis A, or more generally, in a longitudinal direction L within a plane containing the nacelle axis A. The rows of blades 50 preferably extend in a lateral direction perpendicular to the aforementioned direction, such as in a circumferential direction C.

[0056] Thus, when switching from the direct jet configuration ( FIG. 3 ) to the reverse jet configuration ( FIG. 4 ), each shroud 20 moves rearward and exposes the inner and outer sides of the cascade vanes 22 , thereby exposing the cascade vanes 22 to the secondary flow SF on the inner side and to the external environment on the outer side. Furthermore, in the example shown, the displacement of each shroud 20 causes the reverse baffle 32 to deploy in the annular passage 30 of the secondary flow. Consequently, the airflow is deflected toward the cascade vanes 22 by the baffle 32 and flows through the cells 54 while being deflected forward by the blades 50 of the cascade vanes 22 , thereby generating reverse thrust.

[0057] The sliding of each cowl 20 between its forward and rearward positions is typically achieved by cylinders (not shown) distributed around the axis A of the nacelle and attached at the front to a fixed part of the nacelle (e.g. the front frame 44) and at the rear to the cowl 20, for example by suitable fittings.

[0058] 5 , a blade 50 of a prior art cascade (e.g., the cascade 22 shown in FIGS. 3 and 4 ) has a camber angle α, defined as the angle between tangents T1 and T2 to the camber line 60 of the blade, with the tangents T1 and T2 being defined at the leading edge 62 and the trailing edge 64 of the blade, respectively. This camber angle α reflects a more or less pronounced curvature of the blade, on which the aerodynamic performance of the cascade 22 depends. Typical manufacturing processes for blades impose a minimum value on the angle α, thereby limiting the achievable curvature of the blade.

[0059] As described above, in order to improve the aerodynamic performance of the thrust reverser cascade and make the cascade easy to manufacture and low in cost, the present invention provides a cascade 22 having a plurality of stacked blade stages 22A, 22B, such as Figure 6 shown.

[0060] Similar to the cascade blades according to the above-mentioned prior art, the present cascade blade 22 includes a front flange 40 and a rear flange 42, and each blade stage 22A, 22B includes a plurality of rows of blades 50A, 50B shaped to deflect the airflow, for example, toward the front side of the cascade blade, and spars 52A, 52B, for example, in the form of straight beams with a rectangular cross-section. The plurality of rows of blades 50A, 50B intersect the spars 52A, 52B within each blade stage 22A, 22B, so as to form, together with the spars 52A, 52B, a respective honeycomb structure 53A, 53B, which defines a plurality of cells 54A, 54B, each of which forms a passage for the airflow through the cascade blade 22.

[0061] The leading flange 40 and the trailing flange 42 are arranged at the leading and trailing ends of the cascade blade 22, respectively, and are each rigidly connected to the spars 52A, 52B of at least one of the blade stages, or at least to some of the spars 52A, 52B of at least one of the blade stages. Figure 6 In the example shown, the spar 52A of the inner blade stage 22A rigidly connects the leading flange 40 to the trailing flange 42 .

[0062] In order to provide the cascade with the required stability with respect to vibrations and aerodynamic disturbances during operation, the invention further provides that the respective honeycomb structures 53A, 53B of the blade stages 22A, 22B are connected to one another independently of the flanges 40 and 42 by means of a connecting device 66, which is configured to prevent relative displacement between the honeycomb structures 53A, 53B. By "independent of the flanges", it should be understood that the connecting device 66 acts directly on the spar and / or blades of the blade stages 22A, 22B. In other words, the load path in the connecting device 66 does not pass through the flanges 40 and 42. The connecting device 66 is preferably configured to rigidly connect the honeycomb structures 53A, 53B independently of the flanges 40 and 42, as will become more apparent hereinafter.

[0063] The connecting means 66 can be configured to act in the entire area facing the honeycomb structures 53A, 53B, i.e. along the entirety of each row of blades 50A, 50B and each spar 52A, 52B, or only in one or more sections of this facing area, i.e. along one or more sections of one or more spars or only along certain spars and / or along one or more sections of one or more rows of blades or only along certain rows of blades, as will become more apparent hereinafter.

[0064] Reference Figure 7 The spar 52A of the first blade stage 22A has a first extension 70A, while the spar 52B of the second blade stage 22B has a second extension 70B. The first extension 70A and the second extension 70B form the aforementioned connection device 66, or in the example shown, are part of the aforementioned connection device 66. More specifically, the first extension 70A and the second extension 70B are arranged facing each other so as to form a mutual stop that resists relative displacement between the honeycomb structures 53A, 53B, for example relative displacement along the circumferential direction C, or more generally relative displacement along a direction perpendicular to the longitudinal direction L of the spars 52A, 52B.

[0065] To this end, the first extension 70A extends in the direction of the second blade stage 22B, whereas the second extension 70B extends in the direction of the first blade stage 22 A. Thus, the first extension 70A and the second extension 70B extend in the radial direction R, for example.

[0066] Of course, in the embodiment in which the extensions 70A, 70B are in the form of simple stops, the stops act only with respect to the relative displacement between the honeycomb structures 53A, 53B in a given direction. Figure 7 , the displacement that is prevented is the displacement of the lower honeycomb structure 53A oriented in the counterclockwise direction (i.e., in the direction opposite to the direction of the arrow indicating the circumferential direction C) when viewed from the rear, and the displacement of the upper honeycomb structure 53B oriented in the clockwise direction (i.e., in the direction of the arrow indicating the circumferential direction C) when viewed from the rear.

[0067] In order to prevent at least one additional degree of freedom, the first extension 70A and the second extension 70B advantageously comprise mutual engagement means, hereinafter referred to as mutual spar engagement means 72, designed to resist relative displacements of the alveolar structures 53A, 53B, preferably in one or more directions different from the directions in which the extensions 70A, 70B act. Thus, in the example shown, the mutual spar engagement means 72 are designed to resist relative displacements of the alveolar structures 53A, 53B in the radial direction R.

[0068] These mutual spar engagement means 72 define, for example, a sliding connection having an axis 74 parallel to the longitudinal direction L of the spar. To this end, the mutual spar engagement means 72 comprise, for example, a longitudinal groove 76 formed in the second extension 70B, and a longitudinal tongue 78 formed so as to protrude from the first extension 70A, for example in the circumferential direction C, and to be housed—preferably embedded—in the groove 76.

[0069] In this case, extensions 70A and 70B are advantageously flexible to allow groove 76 and tongue 78 to engage each other with a resilient snap fastening. Alternatively, in some embodiments, stages 22A and 22B are assembled by sliding tongue 78 within groove 76 along groove 76.

[0070] exist Figure 7 In the example shown, all spars 52A and 52B comprise respective extensions 70A, 70B provided with mutual spar engagement means 72 .

[0071] In such Figure 8 In the other illustrated embodiments, only some of the spars 52A and some of the spars 52B include respective extensions 70A, 70B.

[0072] Furthermore, the extensions 70A, 70B may extend over the entire length of the spar 52A, 52B in question, or only along a portion of that length. Thus, as desired, provision may be made for the extensions 70A, 70B to extend only along one or more sections of each spar 52A, 52B or some of the spars 52A, 52B, excluding one or more other sections of said spar.

[0073] Reference Figure 9 and Figure 10 It should be understood that, in addition, different relative arrangements of the respective vanes 50A, 50B of the honeycomb structures 53A, 53B are possible within the scope of the present invention.

[0074] Therefore, in Figure 9 In the example shown, the row of blades 50A is offset relative to the row of blades 50B by a distance dR1 along the direction from the leading edge 40 to the trailing edge 42 (i.e., the longitudinal direction L of the spar, which in the example shown corresponds to the axial direction X). Furthermore, in this example, the row of blades 50A is completely radially offset relative to the row of blades 50B. In other words, there is no radial overlap between the row of blades 50A and the row of blades 50B.

[0075] exist Figure 10In the example of FIG, the row of blades 50A is slightly axially offset from the row of blades 50B and has a radial overlap dR1 relative to the row of blades 50B. Thus, it will be appreciated that the blades 50A of the first blade stage 22A include respective trailing edges 64A that extend beyond respective leading edges 62B of the blades 50B of the second blade stage 22B in a direction from the first blade stage 22A to the second blade stage 22B.

[0076] In both cases, each blade 50A of one blade stage 22A forms a blade pair with a corresponding blade 50B of the other blade stage 22B (i.e., the closest blade) that provides airflow redirection performance comparable to that of a single blade having substantially the same bend angle. Such a blade pair has the advantage of being simpler and more economical to manufacture than a single blade having the same characteristics.

[0077] Such an offset between the blade rows of different blade stages may improve the aerodynamic performance of the cascade due to a favorable redistribution of the velocity and pressure profiles in the different stages.

[0078] Furthermore, the spacing εA between consecutive rows of blades within one of the blade stages 22A may be the same as or different from the spacing εB between consecutive rows of blades within another of the blade stages 22B.

[0079] Now refer to Figures 11A-11B In another depicted example, the row of blades 50B may extend in line with the row of blades 50A.

[0080] In this case, the connecting means 66 may advantageously comprise mutual blade engagement means 80 formed on the blades 50A of the first blade stage 22A and the blades 50B of the second blade stage 22B and cooperating to prevent relative displacement of these blades. For illustrative purposes, these means are shown in FIG. Figure 11A In the non-engaged state, Figure 11B The engaged state is shown in the middle.

[0081] These mutual blade engagement means 80 comprise, for example, a tongue 82 fixed to the blade 50A of one of the blade stages 22A and extending in the direction of the other blade stage 22B, and a groove 84 formed in the blade 50B of the other blade stage 22B and in which the tongue 82 is received—preferably embedded—into.

[0082] The tongues 82 of the blades 50A thus extend in the direction of the blades 50B of the other stage. These tongues 82 are oriented, for example, along the radial direction R, which is Figures 11A-11BAs shown, or the tongues are at least oriented along a direction having a radial component (eg a direction D that is radially outwardly inclined in the upstream direction), as shown Figure 11C As in the example shown.

[0083] Consequently, the blades 50A, 50B provided with such a device are prevented from moving relative to each other in the longitudinal direction L of the spar (corresponding in this case to the axial direction X).

[0084] like Figures 11A-11C As shown, the shape of the blades 50A, 50B is such that the blade pair formed by the blades 50A and 50B, assembled or connected to each other by respective mutual blade engagement means 80, substantially defines an aerodynamic profile shape, ie a shape equivalent to that of a conventional blade.

[0085] In other embodiments, the inter-blade engagement arrangement 80 may be configured to prevent displacement in the circumferential direction C or in both directions L and C, or even in the radial direction R or more generally in the direction from the first blade stage 22A to the second blade stage 22B.

[0086] In addition, Figure 12 In the example shown, all rows of blades 50A, 50B are provided with mutual blade engagement means 80, while Figure 13 In the example shown, only some rows of blades 50A, 50B are provided with mutual blade engagement means 80 .

[0087] More specifically, in Figure 13 In the illustrated example, one of the stages 22B has rows of blades 50B that face some rows R1 of blades 50A in the other stage 22A, but not some other rows R2 of blades 50A in the other stage 22A. Consequently, the spacing εA between consecutive rows of blades in the inner stage 22A is smaller than the spacing εB between consecutive rows of blades in the outer blade stage 22B. The aerodynamic profile defined by the blade pairs formed by blades 50A and blades 50B in row R1 can have a more pronounced curvature and / or a greater height than the curvature and / or height of blades 50B in row R2, which can contribute to the aerodynamic performance of the cascade. Furthermore, the relatively large spacing εB between the blades in the aforementioned pairs can help facilitate the manufacture of the cascade, particularly in the case of molded cascades.

[0088] In another configuration, the inner stage 22A and the outer stage 22B are configured relative to Figure 13 The configuration would be the opposite and would provide the same advantages in terms of cascade manufacturing while enabling increased air flow through the cascade due to the reduction in the number of leading edges in the cascade inlet plane, thereby reducing air flow obstruction.

[0089] Similar to the description above regarding the extensions 70A, 70B, the mutual blade engagement means 80 may be defined over the entire length of the row of blades 50A, 50B in question, or only along a portion of that length. Thus, as required, it may be provided that these means extend only along one or more sections of each row of blades or some of the rows of blades, excluding one or more other sections of said rows of blades.

[0090] Figure 14A and 14B Summarizing this principle schematically, in one specific example, the region is shown where connection means 66 are located along sections of rows of blades 50A, 50B and along sections of spars 52A, 52B.

[0091] exist Figure 15 In another example shown, the row of blades 50A (only one of which is shown in isolation) is slightly axially offset from the row of blades 50B (one of which is also visible) and has a radial overlap relative to the row of blades 50B, as shown. Figure 10 Furthermore, the connection means 66 comprises mutual blade engagement means 90 provided in the form of cylindrical hinge supports. For a given pair of blades 50A and 50B, as in Figure 15 As seen in FIG, such a support is defined, for example, by support elements 90A, 90B formed on blades 50A and 50B, respectively, together defining a channel 92 therebetween to allow airflow to circulate between the blades. When the spacing between consecutive spars is relatively large, such a support may advantageously be arranged at a distance from the spars to help strengthen the cascade.

[0092] Figure 16A A two-stage blade cascade 22A, 22B (generally similar to Figure 6 ), which blade cascade is located within a thrust reverser 19 also similar to FIG. 3 .

[0093] In this particular example, the honeycomb structure 53B of the outer stage 22B has a smaller axial extent than the honeycomb structure 53A of the inner stage 22A, which allows the aft end 93B of the honeycomb structure 53B of the outer stage 22B to be offset forward relative to the aft end 93A of the honeycomb structure 53A of the inner blade stage 22A, preferably by a distance corresponding to several rows of blades 50A of the inner stage 22A.

[0094] Thus, the space available for the cascade vanes 22 can be fully utilized since, taking into account the aerodynamic profile of the cowl 20 and, more generally, of the aft section 18 of the nacelle, said space generally decreases towards the rear.

[0095] For this reason, the outer stage 22B therefore includes a smaller number of rows of blades 50B than the inner stage 22A. In the example shown, the rows of blades 50B of the outer stage 22B are still arranged opposite to the consecutive rows of blades 50A of the inner stage 22A, for example, the rows of blades 50A forming the front end portion 94 of the inner stage 22A. Thus, the rear end portion 96 of this stage 22A is formed by the other rows of blades, the trailing edges 64 of which directly face the outer wall 26 of the shroud 20.

[0096] Furthermore, in the example shown, all or part of the connection means 66, such as extensions, such as the above-mentioned extensions 70A, 70B, are defined along a section corresponding to substantially all of the spars 52B of the outer stage 22B. These connection means 66 are arranged as follows: Figure 14A 、 14B The rectangular form shown is very schematically shown in Figure 16A middle.

[0097] It should also be noted that in this example, only the spars 52A of the inner stage 22A, and not the spars 52B of the outer stage 22B, are connected to the leading and trailing flanges 40, 42. Thus, the connecting means 66 provide the entire connection between the outer and inner stages 22B, 22A. This example illustrates the flexibility that the present invention provides in terms of cascade design. In fact, once the blade stages 22A, 22B are connected to each other by their respective honeycomb structures 53A, 53B, it is not necessary to connect each blade stage to the leading and trailing flanges 40, 42. Blade stages (e.g. Figure 16A The outer stage 22B) can therefore extend away from one or both of the flanges 40 and 42.

[0098] Figure 16B A variation is shown in which the axial extension of the honeycomb structure 53B of the outer stage 22B is greater than the axial extension of the honeycomb structure 53A of the inner stage 22A, so that the aft end 93B of the honeycomb structure 53B of the outer stage 22B is offset rearward relative to the aft end 93A of the honeycomb structure 53A of the inner stage 22A. Furthermore, in this example, one of the flanges (e.g., the forward flange 40) is connected to the inner stage 22A, while the other flange (e.g., the aft flange 42) is connected to the outer stage 22B.

[0099] In other embodiments, the cascade 22 according to the present invention may include more than two blade stages, for example Figure 17 Three blade stages are shown schematically.

[0100] Furthermore, the spars of one or more blade stages of the cascade may have an inclination in the circumferential direction C, rather than being oriented only radially as in the above examples. Thus, the blade stages in question trigger a circumferential deflection of the air flow.

[0101] Therefore, in Figure 17 In the embodiment shown, the inner and middle stages 22A, 22C have spars 52A, 52C oriented along the radial direction R, while the outer stage 22B of the cascade has spars 52B oriented along an angle θ relative to the radial direction R in a transverse plane RC.

[0102] As described above, the connecting means 66 may be or include means for fixing the respective honeycomb structures 53A, 53B of the blade stage so as to locally prevent any relative displacement between the structures. For example, such fixing means may take the form of welding or adhesive. In the above example, such fixing means are particularly configured to attach together the extensions 70A and 70B (particularly the tongue 78 and the groove 76), and the tongue 82 and the groove 84, where applicable.

[0103] In view of the above examples, it will therefore be understood that the present invention makes it possible to improve the performance of thrust reverser cascades while providing compatibility with inexpensive manufacturing processes and avoiding or at least limiting the problems of vibration instability and aerodynamic disturbances encountered with two-stage cascades according to the prior art, thanks to the connection between the honeycomb structures.

[0104] Furthermore, the invention allows the configuration of the cascade to be optimized according to the available space, allowing, where applicable, the superposition of blade stages of different lengths in the axial direction X. During the design phase of the nacelle, the options provided by the invention make it possible, for example, to increase the aerodynamic performance of the thrust reverser for a given nacelle length, or to reduce the nacelle length, and thus the nacelle mass, while maintaining consistent aerodynamic performance of the thrust reverser. The invention finds particular application by reducing the number of blade rows in the reverser cascade and compensating this reduction by increasing the number of blade stages, wherein increasing the number of blade stages makes it possible to increase the height of all or some of the remaining blades of the cascade.

Claims

1. A cascade (22) of a cascade thrust reverser (19) for a nacelle of an aircraft propulsion assembly, the cascade comprising: at least a first blade stage (22A) and a second blade stage (22B), said first and second blade stages being superimposed and each comprising a plurality of rows of blades (50A, 50B) shaped to deflect the air flow, and spars (52A, 52B) intersecting the blades within each blade stage so as to form with the blades respective honeycomb structures (53A, 53B) for the passage of the air flow; a front flange (40) and a rear flange (42), which are arranged at the front and rear ends of the cascade (22), respectively, and are each rigidly connected to all or some of the spars (52A, 52B) of at least one of the blade stages; It is characterized in that the corresponding honeycomb structures (53A, 53B) of the blade stage are connected to each other independently of the front flange (40) and the rear flange (42) by connecting devices (66), and the connecting devices are configured to prevent relative displacement between the honeycomb structures (53A, 53B).

2. The cascade according to claim 1, wherein: The connecting device (66) comprises a first extension (70A) of the spar (52A) of the first blade stage and a second extension (70B) of the spar (52B) of the second blade stage, each of the first extensions (70A) being arranged facing a corresponding second extension (70B) such that the first extension and the second extension form a mutual stop that counteracts relative displacement between the honeycomb structures (53A, 53B).

3. The cascade according to claim 2, wherein: The connecting means (66) comprises mutual spar engagement means (72) formed on the first extension (70A) and the second extension (70B) and cooperating to prevent relative displacement between the honeycomb structures (53A, 53B).

4. The cascade according to any one of claims 1 to 3, wherein: The connecting means (66) comprises mutual blade engaging means (80; 90) formed on the blades (50A) of the first blade stage and the blades (50B) of the second blade stage and cooperating with each other to prevent relative displacement between the honeycomb structures (53A, 53B).

5. The cascade according to any one of claims 1 to 4, wherein: The connecting means (66) comprise means for fixing the respective honeycomb structures (53A, 53B) of the blade stages.

6. The cascade according to any one of claims 1 to 5, wherein: The first blade stage (22A) is an inner blade stage, the second blade stage (22B) is an outer blade stage arranged on the inner blade stage, and at least one of the front end and the rear end (93B) of the outer blade stage is offset from the corresponding front end or rear end (93A) of the inner blade stage along the direction (X) from the front flange (40) to the rear flange (42).

7. The cascade according to any one of claims 1 to 6, wherein: The blades (50B) of the second blade stage (22B) extend in line with the corresponding blades (50A) of the first blade stage (22A).

8. The cascade according to any one of claims 1 to 6, wherein the blades (50B) of the second blade stage (22B) are offset relative to the corresponding blades (50A) of the first blade stage (22A) according to a direction from the leading flange (40) to the trailing flange (42).

9. The cascade according to any one of claims 1 to 8, wherein: The blades (50A) of the first blade stage (22A) include respective trailing edges (64) extending beyond respective leading edges (62) of the blades (50B) of the second blade stage (22B) in a direction (R) from the first blade stage (22A) to the second blade stage (22B).

10. A cascade thrust reverser (19) for an aircraft propulsion assembly nacelle, comprising at least one cascade vane according to any one of claims 1 to 9.

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