FR2944328A1 - Stator part for e.g. turbojet engine of aircraft, has braze extended towards back on both sides of head of vane while being stopped in front of trailing edge of vane, where trailing edge is arranged in rear part of transversal perforation - Google Patents

Abstract

The part has stator vanes (4), and an exterior shroud that is provided with transversal perforations (22), where each perforation receives a head (24) of one of the stator vanes. The head is fixed to the exterior shroud by a braze (26) arranged in the corresponding perforation. The braze is extended towards the back on both sides of the head of the vane while being stopped in front of a trailing edge (30) of the vane, where the trailing edge is arranged in a rear part (22a) i.e. oblong hole, of the perforation.

Description

STATOR FOR AIRCRAFT TURBOMOTOR COMPRISING AN EXTERNAL VIROLE HAVING A MECHANICAL AUBILLATION HOLES DESCRIPTION

The present invention relates generally to an aircraft turbine engine, preferably of the turbojet or turboprop type. More particularly, the invention relates to a stator portion for a module of such a turbine engine, this part comprising a plurality of stator vanes as well as an outer shell intended to delimit radially outwardly a primary flow passing through this turbine engine . To maintain the stator vanes extending radially inwardly from the outer shell, the latter usually comprises a plurality of circumferentially spaced, and through-formed, openings, each of which is formed in the outer shell. thickness of the ferrule, generally metallic. Each aperture houses the head of one of the stator vanes, this head being fixed to the outer ferrule by means of a solder made in the aperture concerned, extending continuously all around the head of the stator. 'dawn. This design, however widely spread in the field, has the disadvantage that cracks frequently appear at the trailing edge of the blade head. The explanation of this phenomenon harmful to the life of the stator blades lies in the presence of very large static charges at the trailing edge of the blade head, which result in a dynamic margin too low, favoring the appearance of creeks. To cope with this drawback, it has been proposed to thicken the outer shell, so as to limit its deformations in operation, and thus to reduce the stresses applied to the trailing edges of the blade heads. Nevertheless, this solution is extremely penalizing in terms of cost and mass.

It has also been proposed to thicken the blades, but this inevitably leads to a disturbance of the air flow therethrough, with a consequent negative impact on the overall performance of the turbine engine.

The invention therefore aims to at least partially overcome the problem mentioned above, relating to the achievements of the prior art. To do this, the subject of the invention is a stator portion for an aircraft turbine engine module comprising a plurality of stator vanes, as well as an outer ferrule provided with a plurality of through openings each receiving the head of the turbine engine. one of said vanes, attached to the outer shell by a solder arranged in its associated aperture.

According to the invention, for at least one of said stator vanes, said solder extends rearwardly on either side of the blade head, stopping before a trailing edge portion of it, arranged in a rear portion of said through aperture.

In other words, the invention is remarkable in that the solder no longer extends all around the blade head in the through aperture, but leaves free a trailing edge portion. This overall makes it possible to reduce the static charges at this specific location of the blade head, to release the dynamic margin, and thus to strongly delay the appearance of cracks. The rear part of the through aperture therefore fulfills the function of mechanical unloading of the blade.

The solution provided is very satisfactory in that it generates no penalty in terms of mass, cost or performance. In addition, this results in an increased blade life, advantageously involving the periods of visit / maintenance may be further spaced over time. Naturally, this specificity can be applied to all the stator vanes or only some of them, depending on the needs and constraints encountered. By way of indicative example, in the case where the stator part according to the invention is sectorised, it has been found advisable to make the trailing edge portion free only on the two blades located at the ends of each angular portion. Nevertheless, any other configuration can be envisaged without departing from the scope of the invention. Preferably, said rear portion of said through aperture, that is to say that without solder, takes the form of a through oblong hole whose length is oriented substantially in a direction orthogonal locally to the skeleton of dawn head . This particular form, allowing to release a very satisfactory dynamic margin, can be easily obtained during the realization of the opening in the ferrule.

Preferably, said trailing edge portion extends along a length (1), taken along the blade head skeleton, between 0.03 and 0.05 times the total length of said blade head. along this skeleton. This feature allows to further reduce the risk of occurrence of cracks on the stator blade heads. For example, filling material may fill the space between said trailing edge portion and the rear portion of said through aperture. The main function sought by the introduction of this material is the sealing of the air flow passing through the stator vanes, but also limits thermal losses. Of course, this material is chosen so as to exert only a null or negligible mechanical stress on the trailing edge portion of the blade head. Thus, either the rear portion of the aperture remains hollow, or it is filled by said filling material which has a mechanical hold of the trailing edge portion lower than that which could be provided by the solder, this mechanical retention to the ferrule by the filling material nevertheless remaining, as mentioned above, zero or negligible. For example, it may be a filler material of the abradable type, for example a temperature-resistant silicone rubber compound 5 loaded with beads. Preferably, the stator portion further comprises an inner ferrule provided with a plurality of through apertures each receiving the foot of one of said vanes, fixed to the inner ferrule by a solder arranged in its associated aperture. Here also, it is possible to ensure that for at least one of said stator vanes, said solder extends rearwardly on either side of the blade root stopping before a portion of trailing edge thereof, arranged in a rear portion of said through aperture. This allows, in a manner similar to that indicated above, to limit the appearance of cracks at this specific location of the blade also mechanically constrained. The stator portion forms a substantially annular structure, with the outer shell formed for example continuously, preferably in one piece.

Alternatively, the outer shell is formed by the implementation end to end of angular sectors of outer ring. These two alternatives are also possible for the inner ferrule.

The subject of the invention is also an aircraft turbine engine module comprising at least one stator part as described above, the module preferably being a compressor, preferably a low-pressure one, but which can alternatively be a turbine, without leaving of the scope of the invention.

Finally, the invention relates to a turbine engine for aircraft comprising at least one module as described above, the turbine engine being preferably a turbojet. Other advantages and features of the invention will become apparent from the detailed non-limiting description below. This description will be made with reference to the appended drawings among which; Figure 1 shows a partial schematic half-sectional view of a low-pressure turbine engine compressor according to a preferred embodiment of the present invention; FIG. 2 is a front view of an angular portion assembly of the stator portion of the compressor shown in FIG. 1; - Figure 3 shows a partial top view of the assembly shown in Figure 2; - Figure 4 shows a partial sectional view taken along the skeleton of one of the vanes 25 of the assembly shown in Figures 2 and 3; and FIG. 5 represents a partial sectional view taken along the line VV of FIG. 3. Referring firstly to FIG. 1, part of a low pressure compressor 1 for a turbine engine can be seen, according to a preferred embodiment of the present invention. In a known manner, the compressor has, alternately in an axial direction parallel to the axis 2 of the compressor, stator vanes 4 and rotor blades 6. The stator vanes 4, distributed circumferentially about the axis 2 , are integrated in a stator portion 13 also object of the invention, further comprising an inner ferrule 8 of radially inner delimitation of a primary annular flow 10 passing through the turbine engine, this ferrule 8 carrying the feet of the blades 4 which the through. It also contains an outer ring 16 of external radial delineation of the primary annular flow 10, which carries the heads of the vanes 4 which pass therethrough.

In this regard, it is noted that the stator portion also comprises known additional elements reported on the ferrule 8, such as a radially inner abradable coating 15 forming annular sealing track, contacted by a sealing device 12 carried by the rotor stage 14 carrying the rotating vanes 6 and arranged downstream of the stator portion concerned. As mentioned above, the rotating sealing device 12 is in a known manner of labyrinth or wipe type.

In the preferred embodiment described, the stator portion forms a substantially annular structure of axis 2, and is preferably composed by the establishment of a plurality of assemblies 20 such as that shown in FIG. angular or circumferential portion of this stator portion. As shown in this figure, each set 20 comprises an inner ferrule angular sector 8a carrying a plurality of stator vanes 4, the inner ferrule 8 resulting from the end-to-end placement of all sectors 8a.

In the segmented configuration shown, the angular sectors 8a (only one visible in FIG. 2) jointly forming the shell 8 are therefore preferably devoid of direct rigid mechanical connections connecting them to each other, their adjacent ends being in fact simply placed in position. the segmentation operated for the shell 8 is likewise adopted, in the same way, for the abradable radially inner lining 15 forming an annular sealing strip, only an angular sector 15a of this coating thus being fixedly supported by the ferrule sector 8a. In a similar manner, the assembly 20 comprises an angular sector of outer shell 16a carrying the plurality of stator vanes 4, the outer shell 16 thus resulting from the establishment of all the sectors 16a. Here too, the angular sectors 16a (only one visible in FIG. 2) jointly forming the shell 16 are therefore preferably free of direct rigid mechanical connections connecting them to each other, their adjacent ends being in fact simply placed opposite each other. others, with or without play

By being carried out by the end-to-end insertion of several of these angular portions 20, the stator part 13 thus adopts a so-called segmented design, as opposed to a so-called continuous design, in which the ferrules 8 and 16 are each unique and continuous over 360 °. An alternative could also be to provide only one of the two ferrules 8 and 16 is unique and continuous 360 °, and the other shell sectorized. As an indication, it is noted that the number of assemblies / sectors 20 to form the annular stator portion 13 above may be between 6 and 14, each set 20 preferably having the same angular / circumferential extent. Referring now to Figure 3, it can be seen that the outer annular sector 16a of outer shell is provided with perforations 22 made in the thickness of this sector 16a, conferring them a character through. For one of the stator vanes 4 shown, namely that of the left in the figure, the head of the blade 24 is received in its associated aperture 22. It is fixed therein rigidly by a solder 26, made in space 28 delimited between the wall of the ajour and the outer surface of the blade head. Nevertheless, one of the peculiarities of the present invention lies in the fact that the solder 26 is not continuous all around the outer surface of the blade head 24, but is stopped so as to leave free a portion of edge leakage 30 of the dawn head. In other words, this portion 30, which is housed in a rear portion 22a of the aperture 22, is not connected to the sector 16a by the solder 26, the latter in fact traveling continuously on the other hand and other of the head 24, by two sections initiated at the front of the trailing edge portion 30 and joining at the leading edge. The solder 26 thus generally takes a continuous form of stretched U filling most of the space 28. As can be seen in FIG. 3, the rear part 22a of the aperture 22, which is not filled by the solder 26, takes the form of a through oblong hole, preferably oriented substantially in a direction orthogonal locally to the skeleton 32 of the blade head 24. In this embodiment, the solder 26 stops indeed at the entrance to the hole oblong through 22a, a portion of which is filled by the trailing edge portion 30 which can deform substantially freely in dynamic, and the other part can remain hollow. Alternatively, this other part could be filled by a filling material essentially for sealing the primary flow, namely to prevent leakage through the hole 22a. Such a filling material is shown schematically by the reference numeral 36 in FIG. 5. It may for example be a filler material of the abradable type, for example a temperature-resistant silicone rubber compound loaded with balls. . Referring now to FIG. 4, the trailing edge portion 30 extends along a length 1, taken along the skeleton 32, between 0.03 and 0.05 times the total length (unreferenced) of the dawn head 24 along this skeleton. 11 If such an oblong hole of mechanical blade discharge 22a is preferably provided only on the two blades located respectively at both ends of the assembly 20, it could nevertheless be otherwise, without departing from the scope of the invention. In addition, such oblong holes mechanical blade discharge could also be provided at the through openings of the inner shell 8, which receive the blade roots.

The assembly would then be identical or similar to that shown in Figures 3 to 5 for the outer shell. Of course, various modifications may be made by those skilled in the art to the invention which has just been described, solely as non-limiting examples.

Claims (10)

REVENDICATIONS1. A stator part (13) for an aircraft turbine engine module comprising a plurality of stator vanes (4), and an outer shroud (16) provided with a plurality of through apertures (22) each receiving the head (24) of one of said blades, fixed to the outer shell by a solder (26) arranged in its associated aperture (22), characterized in that for at least one of said stator vanes (4), said solder (26) extends rearwardly on either side of the blade head (24) stopping before a trailing edge portion (30) thereof, arranged in a rear portion ( 22a) of said through aperture. 2. Stator portion according to claim 1, characterized in that said rear portion (22a) of said through aperture takes the form of a through oblong hole whose length is oriented substantially in a direction orthogonal locally skeleton (32) head dawn (24). Stator portion according to claim 1 or 2, characterized in that said trailing edge portion (30) extends along a length (1), taken along the blade head skeleton, between 0.03 and 0.05 times the total length of said blade head (24) along this skeleton (32). A stator part according to any one of the preceding claims, characterized in that a filling material (36) fills the space between said trailing edge portion (30) and the rear portion (22a) of said through aperture (22). 5. Stator part according to any one of the preceding claims, characterized in that it further comprises an inner shell (8) provided with a plurality of through holes each receiving the foot of one of said vanes, fixed to the inner shell by a solder arranged in its associated opening. 6. Stator part according to any one of the preceding claims, characterized in that it forms a substantially annular structure, with the outer shell (16) continuous. 7. Stator part according to any one of claims 1 to 5, characterized in that it forms a substantially annular structure, with the outer shell (16) formed by the establishment end to end of angular sectors of outer ring (16a). 8. Module (1) of an aircraft turbine engine comprising at least one stator part (13) according to any one of claims 1 to 7.30. 9. Module according to claim 8, characterized in that it is a compressor, preferably low pressure. 10. Aircraft turbine engine comprising at least one module (1) according to claim 8 or claim 9.10