Rotating assembly of an aircraft turbomachine comprising a fan platform with a stop

Abstract

The invention relates to a rotating assembly of an aircraft turbomachine (1), comprising a disc (22) having at least one tooth (25) made at the periphery of the disc (22), and at least one blade platform (30) mounted on a tooth (25) of the disc (22), characterized in that the platform (30) has, at its upstream end (30a), a stop lug (45) having a first portion (45a) extending substantially axially upstream from the upstream end (30a) of the platform (30), and a second portion (45b) extending substantially radially inwards from the upstream end (45aa) of the first portion (45a) to come into contact with an upstream face (27aa) of the tooth (25) on which the second portion (45b) of the stop lug (45) of the platform (30) is mounted. Figure for the summary: Figure 5

Description

Title of the invention: Rotating assembly of an aircraft turbomachine comprising a fan platform with a stop TECHNICAL FIELD OF THE INVENTION

[0001] The present invention relates to the general field of turbomachines, and more specifically to the field of attached fan blade platforms of a turbomachine.

[0002] The invention applies to any type of aeronautical turbomachines, and in particular to aircraft turbomachines such as turbojets and turboprops.

[0003] The invention thus proposes a rotary assembly of an aircraft turbomachine comprising a blade platform provided with a stop tab, an aircraft turbomachine fan comprising such an assembly, as well as an aircraft turbomachine comprising such a fan. STATE OF THE ART

[0004] In a turbomachine, for example a turbojet, the blades, composed of a root and a blade defined by two sides called respectively intrados and extrados, are mounted in groove-shaped housings, called cells or grooves, specially arranged on the periphery of the rotor disks. These grooves can be straight or curvilinear. Inter-blade platforms are generally mounted in the intervals between two adjacent blades.

[0005] In the case of a fan, the inter-blade platforms must perform several functions. From an aerodynamic point of view, the primary function of these platforms is to define the air flow path. In addition, they must also be able to withstand significant forces without deforming and while remaining attached to the disc that carries them.

[0006] It is also imperative that the inter-blade platforms of the fan can be held in position in order to avoid any wear problems due to their movement. In other words, it is necessary to ensure that the platforms cannot move axially and radially.

[0007] In order to satisfy these different requirements, certain configurations have for example been proposed in which the platforms have a first part making it possible to define the air flow path and to ensure the retention of the platform when the engine is rotating, and a second part making it possible to limit the deformations of the first part under the effects of centrifugal forces and to maintain the platform in position when the engine is stopped.

[0008] In existing solutions, the platform can for example take the form of a box with a two-dimensional vein wall retained downstream by a drum and upstream by a shell, the upstream retention by the shell being carried out above the tooth of the fan disc, a flange of the shell blocking the platform axially and radially upstream.

[0009] As examples, international applications WO 2017 / 006054 A1, WO 2018 / 007717 A1 and WO 2020 / 260810 A1, European patent applications EP 1 970 537 A1, EP 3 663 530 A1 and EP 3 677 752 A1, and American patent application US 2019 / 0145283 A1 of the Applicant describe proposed configurations for limiting the movements of the fan inter-blade platforms.

[0010] However, it happens that the locking in position of the fan inter-blade platforms is insufficient and / or incorrect so that they can move axially and / or radially, then causing wear by friction at the interfaces between the platform and the upstream shell and / or the platform and the downstream drum.

[0011] There is thus a need to ensure that the fan inter-blade platforms are held in position. In particular, there is a need to prevent any axial and / or radial movement of the platforms likely to cause undesirable wear at the interfaces with the inter-blade platforms. Statement of the invention

[0012] The invention aims to at least partially remedy the needs mentioned above and the drawbacks relating to the embodiments of the prior art.

[0013] The invention thus relates, according to one of its aspects, to a rotary assembly of an aircraft turbomachine, comprising a disk, in particular a fan disk, provided with at least one tooth made on the periphery of the disk, and at least one blade platform mounted on a tooth of the disk,

[0014] characterized in that the platform comprises, at its upstream end, a stop tab comprising a first portion extending substantially axially upstream from the upstream end of the platform, and a second portion extending substantially radially inwards from the upstream end of the first portion to come into contact with an upstream face of the tooth on which the second portion of the stop tab of the platform is mounted.

[0015] Thanks to the invention, it is possible to ensure that the platform is held in position by means of the presence of a stop tab so as to prevent or minimize wear caused by movements of the platform.

[0016] The rotating assembly of an aircraft turbomachine according to the invention may further comprise one or more of the following characteristics taken in isolation or in any possible technical combination.

[0017] The platform may have a box shape comprising an outer wall of vein, an inner bottom wall and two side walls extending radially between the inner bottom wall and the outer vein wall, the stop tab extending from the inner bottom wall.

[0018] In addition, the stop tab can be made in one piece with the platform, in particular with the internal bottom wall.

[0019] The rotating assembly of the aircraft turbomachine may also comprise an upstream shell fixed to the disk and axially blocking the platform upstream, the stop tab being located between the upstream shell and the upstream face of the tooth.

[0020] The first portion of the stop tab may comprise an outer surface intended to come into contact with the upstream ferrule.

[0021] The tooth may comprise a notch formed in the axial extension of the tooth upstream and defining an upstream face of the notch of the tooth. In addition, the second portion of the stop tab may be housed in the notch in contact with the upstream face of the notch of the tooth.

[0022] The largest axial dimension of the second portion of the stop tab may advantageously correspond substantially to the largest axial dimension of the notch.

[0023] Furthermore, the tooth may comprise two lateral notches formed in the axial extension of the tooth upstream, in particular on either side of a central portion of the tooth intended to come into contact with an upstream ferrule. In addition, the platform may comprise, at its upstream end, two stop tabs, each being housed in a notch by means of its second portion.

[0024] The second portion of the stop tab may also be a fixing flange intended to be sandwiched between the upstream face of the tooth and an upstream ferrule. The rotating assembly of the aircraft turbomachine may then comprise fixing means, in particular of the screw / nut type, for fixing the second portion to the tooth.

[0025] Furthermore, the invention also relates, according to another of its aspects, to an aircraft turbomachine fan, comprising a rotary assembly of an aircraft turbomachine as defined above comprising a fan disk provided with a plurality of teeth, a plurality of blades whose roots are mounted in grooves, each formed between two adjacent teeth, at the periphery of the disk, and a plurality of platforms mounted on the teeth of the disk.

[0026] Furthermore, the invention also relates, according to another of its aspects, to an aircraft turbomachine comprising a fan as defined above. BRIEF DESCRIPTION OF THE FIGURES

[0027] Other advantages, aims and particular characteristics of the invention will emerge from the following non-limiting description of at least one mode of implementation of the present invention, with reference to the appended figures, in which: • [Fig.l] schematically represents in section an example of a turbomachine in accordance with the invention, • [Fig.2] is a schematic view along direction II of the fan of the turbomachine of [Fig.l], • [Fig.3] is a partial schematic view in axial section of an example of an aircraft turbomachine fan, such as that of [Fig.l], • [Fig.4] is a partial schematic perspective view of the fan disc of the fan of [Fig.3], seen from upstream, • [Fig.5] is a partial schematic perspective view of an example of an aircraft turbomachine fan according to the invention, seen from upstream, • [Fig.6] is a partial schematic view in axial section and in perspective of the turbomachine fan of [Fig.5], and • [Fig.7] illustrates schematically and partially in axial section an alternative embodiment of an example of an aircraft turbomachine fan in accordance with the invention.

[0028] Throughout these figures, identical references may designate identical or similar elements.

[0029] Furthermore, the different parts shown in the figures are not necessarily shown on a uniform scale, in order to make the figures more readable. DETAILED DESCRIPTION OF THE INVENTION

[0030] Throughout the description, given as a non-limiting example of embodiment, it is noted that the terms upstream and downstream are to be considered with respect to a main direction F of normal flow of the gases (from upstream to downstream) for a turbomachine 1. Furthermore, the axis X of the turbomachine 1 is called the axis of radial symmetry of the turbomachine 1. The axial direction of the turbomachine 1 corresponds to the axis of rotation X of the turbomachine 1. A radial direction of the turbomachine 1 is a direction perpendicular to the axis X of the turbomachine 1.

[0031] Furthermore, unless otherwise specified, the adjectives and adverbs axial, radial, axially and radially are used with reference to the aforementioned axial and radial directions. Furthermore, unless otherwise specified, the terms inner and outer are used with reference to a radial direction so that the inner part of an element is closer to the X axis of the turbomachine 1 than the outer part of the same element.

[0032] With reference to [Fig.l], a turbomachine 1 according to the invention of the double-flow turbojet type centered on an axis X is shown schematically in longitudinal section. It comprises from upstream to downstream: a fan 2, a low-pressure compressor 3, a high-pressure compressor 4, a combustion chamber 5, a high pressure turbine 6 and a low pressure turbine 7.

[0033] [Fig.2] represents a schematic view of the blower 2 of [Fig.l] according to the direction II. The fan 2 comprises a fan disc 22 in which a plurality of grooves 24, or cells, is formed at its outer periphery. These grooves 24 are for example rectilinear and extend axially from upstream to downstream along the entire length of the disc 22. They are also regularly distributed all around the axis X of the disc 22. In this way, each groove 24 defines with its neighbor a tooth 25 which also extends axially from upstream to downstream along the entire length of the disc 22. Equivalently, a groove 24 is delimited between two neighboring teeth 25.

[0034] The fan 2 further comprises a plurality of blades 26 of curvilinear profile, only four blades 26 being shown in [Fig. 2]. Each blade 26 has a root 26a which is mounted in a respective groove 24 of the fan disc 22. For this purpose, the root 26a may have a fir tree or dovetail shape adapted to the geometry of the grooves 24.

[0035] In addition, the fan 2 comprises a plurality of attached platforms 30, each platform 30 being mounted in the interval between two neighboring fan blades 26, in the vicinity of the roots 26a thereof, in order to delimit, on the inner side, an annular air inlet vein into the fan 2, the vein being delimited on the outer side by a fan casing.

[0036] Figures 3 and 4 further illustrate an example of a fan 2 comprising locking means for blocking the platforms 30 of the blades 26 in the radial and tangential directions, as described for example in the international application WO 2017 / 006054 A1 of the Applicant. [Fig. 3] thus represents a fan 2 of a turbomachine 1, for example such as that described previously with reference to [Fig. 1], seen in section at the level of a tooth 25 of the fan disc 22, according to the section plane III visible in [Fig. 4].

[0037] The fan 2 comprises a fan disk 22 mounted at the upstream end of a turbine shaft 14 and which carries at its periphery a plurality of blades 26 regularly distributed around the axis X of rotation of the fan 2 and between which are mounted inter-blade platforms 30. Each platform 30 here takes the form of a box with a vein wall 31, forming the upper or outer face of the platform 30 and having an inclined profile, a bottom wall 32 which allows the platform 30 to rest on the tooth 25 of the fan disk 22 when the turbomachine 1 is stopped, and two side walls 33 extending radially between the bottom wall 32 and the vein wall 31.

[0038] The blades 26 are surrounded externally by a nacelle, not shown, which defines with the platforms 30 an annular flow vein for the air flow F entering turbomachine 1.

[0039] The disc 22 comprises at its periphery an alternation of grooves 24 and teeth 25 extending longitudinally over the entire length of the disc 22. The blades 26 comprise at their radially internal ends roots which are engaged axially from upstream in the grooves 24 of the disc 22 and which cooperate by shape connection with these grooves 24 in order to ensure the radial retention of the blades 26 on the disc 22. The blade roots are for example in this case in the shape of a dovetail.

[0040] The fan 2 also comprises a downstream drum 16, or rotating spacer, also called a “booster” drum, “booster” shell or “rotating spacer”, provided with a retaining flange 16a intended to cooperate with an assembly step 31a of the platform 30. The downstream drum 16 is attached and fixed to the downstream face of the fan disk 22 at the level of a bore 25b made in a step downstream of the tooth 25, by means of a fixing 29, in particular of the screw / nut type. This downstream drum 16 cooperates in a sealed manner at its radially external end with an internal casing 13 of the fan 2 which is aligned with the inter-blade platforms 30.

[0041] Furthermore, the fan 2 also comprises an upstream ferrule 42 which is fixed to the fan disc 22 at the level of the lug 25a which axially extends the tooth 25 of the fan disc 22 upstream, by means of fixings 42a, in particular of the screw / nut type. The upstream ferrule 42 is aligned with the inter-blade platforms 30.

[0042] In the embodiment of Figures 3 and 4, the platform 30 comprises at its upstream end a locking ring 34, of trapezoidal section, in which the tab 25a of the tooth 25 of the fan disc 22 is intended to be housed.

[0043] The fan disc 22 is coupled with the turbine shaft 14 of the turbomachine 1 so that in operation, the upstream shroud 42, the fan disc 22, the blades 26 and the downstream drum 16 are driven integrally in rotation by the low-pressure turbine 7. The platforms 30, mounted between the upstream shroud 42 and the downstream drum 16, and axially blocked by these two elements, are thus also driven integrally in rotation.

[0044] With reference to Figures 5 and 6, an exemplary embodiment of a fan 2 comprising a rotating assembly of an aircraft turbomachine 1 according to the invention is shown.

[0045] As previously described, the fan disc 22 comprises a plurality of teeth 25 between which grooves 24 are formed at the periphery of the disc 22. On each tooth 25 is mounted an inter-blade platform 30.

[0046] The platform 30 comprises, at its upstream end 30a, a stop tab 45 which will serve as a stop on the fan disc 22 and which will be blocked by the upstream ferrule 42 before the latter is screwed onto the fan disc 22. in this way, the interaction between the stop tab 45 and the fan disc 22, as well as the interaction between the stop tab 45 and the upstream shroud 42, make it possible to immobilize the fan blade platform 30 26 and thus to prevent or minimize the wear phenomenon caused by the axial and / or radial movements of the platform 30.

[0047] Specifically, the stop tab 45 comprises a first portion 45a which extends substantially axially upstream from the upstream end 30a of the platform 30, and a second portion 45b which extends substantially radially inwards, in other words towards the axis of rotation X, from the upstream end 45aa of the first portion 45a so as to be in contact with an upstream face 27aa of the tooth 25 on which the second portion 45b is mounted.

[0048] In the example shown here in Figures 5 and 6, each tooth 25 has two lateral notches 27 formed in the axial extension of the tooth 25 upstream, on either side of a central portion 25p of the tooth 25 which comes into contact with the upstream ferrule 42.

[0049] Each notch 27 defines an upstream face 27aa and a second portion 45b of a stop tab 45 can thus be received in a notch 27 and be in contact with the upstream face 27aa. Here, each platform 30 comprises a stop tab 45 whose second portion 45b comes into contact with an upstream face 27aa of a notch 27. However, each platform 30 could comprise two stop tabs 45, one for each notch 27.

[0050] The second portion 45b thus defines a radial contact surface intended to come into contact with a corresponding surface of the tooth 25 which is therefore the upstream face 27aa. In addition, the first portion 45a comprises an external surface 45as intended to come into contact with the upstream ferrule 42 after assembly.

[0051] Advantageously, as visible in [Fig.6], the largest axial dimension D45b of the second portion 45b of a stop tab 45 corresponds substantially to the largest axial dimension D27 of a notch 27.

[0052] After assembly of the upstream shell 42 on the fan disk 22, as visible in [Fig. 6], the interaction between the stop tab 45 and the disk 22 on the one hand, and between the stop tab 45 and the upstream shell 42 on the other hand, prevent radial movements as well as axial movements of the platform 30 relative to the disk 22 and the upstream shell 42. Immobilizing the platform 30 relative to the disk 22 and the upstream shell 42 therefore makes it possible to prevent or minimize wear at the interfaces between the platform 30 and the upstream shell 42, and between the platform 30 and the downstream drum 16.

[0053] Furthermore, [Fig.7] illustrates another possibility of producing the stop tab. 45.

[0054] In this example, the second portion 45b of the stop tab 45 is a fixing flange sandwiched between the upstream face 25aa of the tooth 25 and the upstream ferrule 42. Then, fixing means 46, in particular of the screw / nut type, are used to fix together the upstream ferrule 42, the second portion 45b and the tooth 25.

[0055] The second portion 45b in the form of a flange also makes it possible to prevent or minimize wear at the interfaces between platform 30 and upstream shell 42, and between platform 30 and downstream drum 16.

[0056] In the embodiments described above, the platform 30 may have a box shape comprising an outer vein wall 31, an inner bottom wall 32 and two side walls 33 extending radially between the inner bottom wall 32 and the outer vein wall 31. The stop tab(s) 45 may then extend from the inner bottom wall 32. In addition, the stop tab(s) 45 are advantageously made in one piece with the inner bottom wall 32 of the platform 30.

[0057] Of course, the invention is not limited to the embodiments which have just been described. Various modifications can be made thereto by those skilled in the art.

Claims

Claims

1. Rotary assembly of an aircraft turbomachine (1), comprising a disc (22) provided with at least one tooth (25) produced at the periphery of the disc (22), and at least one platform (30) of a blade (26) mounted on a tooth (25) of the disc (22), characterized in that the platform (30) comprises, at its upstream end (30a), a stop tab (45) comprising a first portion (45a) extending substantially axially upstream from the upstream end (30a) of the platform (30), and a second portion (45b) extending substantially radially inward from the upstream end (45aa) of the first portion (45a) to come into contact with an upstream face (25aa, 27aa) of the tooth (25) on which the second portion (45b) of the stop tab (45) of the platform (30) is mounted.

2. An assembly according to claim 1, wherein the platform (30) has a box shape comprising an outer vein wall (31), an inner bottom wall (32) and two side walls (33) extending radially between the inner bottom wall (32) and the outer vein wall (31), the stop tab (45) extending from the inner bottom wall (32).

3. Assembly according to claim 1 or 2, in which the stop tab (45) is made in one piece with the platform (30), in particular with the internal bottom wall (32).

4. Assembly according to one of the preceding claims, in which it comprises an upstream ferrule (42) fixed on the disc (22) and axially blocking the platform (30) upstream, the stop tab (45) being located between the upstream ferrule (42) and the upstream face (27aa) of the tooth (25).

5. Assembly according to any one of the preceding claims, in which the tooth (25) comprises a notch (27) formed in the axial extension of the tooth (25) upstream and defining an upstream face (27aa) of the notch (27) of the tooth (25) and in which the second portion (45b) of the stop tab (45) is housed in the notch (27) in contact with the upstream face (27aa) of the notch (27) of the tooth (25).

6. An assembly according to claim 5, wherein the largest axial dimension (D45b) of the second portion (45b) of the stop tab (45) corresponds substantially to the largest axial dimension (D27) of the notch (27).

7. Assembly according to claim 5 or 6, in which the tooth (25) comprises two lateral notches (27) formed in the axial extension of the tooth (25) upstream, on either side of a central portion (25p) of the tooth (25) intended to come into contact with an upstream ferrule (42), and in which the platform (30) comprises, at its upstream end (30a), two stop tabs (45), each being housed in a notch (27) by means of its second portion (45b).

8. Assembly according to any one of claims 1 to 4, in which the second portion (45b) of the stop tab (45) is a fixing flange intended to be sandwiched between the upstream face (25aa) of the tooth (25) and an upstream ferrule (42) and in which it comprises fixing means (46), in particular of the screw / nut type, for fixing the second portion (45b) to the tooth (25).

9. Fan (2) of an aircraft turbomachine (1), comprising a rotary assembly of an aircraft turbomachine (1) according to any one of the preceding claims comprising a fan disk (2) provided with a plurality of teeth (25), a plurality of blades (16) whose roots (26a) are mounted in grooves (24), each formed between two adjacent teeth (25), at the periphery of the disk (22), and a plurality of platforms (30) mounted on the teeth (25) of the disk (22).

10. Aircraft turbomachine (1) comprising a fan (2) according to claim 9.