EP1855011B1 - Assembly for the engine compressor of an aircraft comprising tilted-head hammer attachment vanes - Google Patents

Description

TECHNICAL AREA

The present invention relates generally to a disk / blade assembly for an aircraft engine compressor, comprising a disk as well as a plurality of hammer link blades mounted on the same disk, and more specifically in a circumferential groove of this one.

Preferably, the application relates to the high-pressure compressor of an aircraft engine such as a turbojet engine or a turboprop engine, and preferably the rear stages of this compressor. However, the invention could also be applied for the low-pressure compressor without departing from the scope of the invention.

The invention also relates to an aircraft engine compressor, high pressure or low pressure, equipped with at least one such disk / blade assembly, and an aircraft engine equipped with at least one such compressor .

STATE OF THE PRIOR ART

The document EP 1 219 782 is considered the document describing the closest state of the art.

From the prior art, a disk / blade assembly for an aircraft engine compressor, comprising a disc as well as a disc, is known. plurality of hammer blade mounted on said disk, wherein each blade successively comprises, in a radially inward direction, a blade, a platform, a stilt, and a blade root provided with a bearing surface upstream located at a leading edge of the blade as well as a downstream bearing surface located at a trailing edge side of the blade.

In addition, the disk is provided with a circumferential groove in which is retained the blade root of each of the vanes, by means of bearing surfaces bearing against this circumferential groove provided for this purpose. This allows to retain the blades in the radial direction outwardly, relative to the disc in which their blade root is housed.

It has been noted in the prior art embodiments that the intensity of the mechanical stresses encountered at the bearing surfaces and the stilt was extremely irregular, obviously involving design problems.

STATEMENT OF THE INVENTION

The object of the invention is therefore to provide a disk / hammer blade assembly that overcomes the aforementioned problem relating to the embodiments of the prior art.

To do this, the subject of the invention is a disk / blade assembly for an aircraft engine compressor, comprising a disk as well as a plurality of hammer blade blades mounted on this disk, each blade comprising successively, in one direction radially inward, a blade having a leading edge and a trailing edge offset circumferentially from the leading edge in a given offset direction, a platform, a stilt, and a blade root provided with a surface an upstream bearing surface located at a leading edge side of the blade as well as a downstream bearing surface situated at a trailing edge side of this blade, the disk being provided with a circumferential groove in which is retained the blade root of each of the plurality of blades through the bearing surfaces bearing against this circumferential groove. According to the invention, for each of the plurality of blades, the downstream bearing surface is circumferentially offset from the upstream bearing surface in the given direction of given shift.

Therefore, the invention advantageously proposes to modify the geometry of the blade roots used until now which was to extend each foot parallel to a central axis of the disk, going from its upstream surface surface to its surface downstream range. Indeed, in the proposed configuration where the downstream bearing surface is circumferentially offset from the upstream bearing surface in the given offset direction corresponding to the offset direction of the trailing edge of the blade relative to the leading edge of the blade. ci, the advantageous consequence lies in the fact that the blade root and its associated stilt follow substantially the profile of the blade. In other words, in top view of a given dawn, the size of the intersection between the blade root and the Pale is therefore greatly increased compared to that encountered in the prior art, where this magnitude remained relatively low because of the low compatibility between the orientation of the foot along the central axis of the disc, and the geometry of the profiled blade.

This then makes it possible to obtain a better homogeneity in the intensity of the mechanical stresses encountered at the level of the bearing surfaces and the stilt, which thus advantageously reduces considerably the design difficulties encountered previously.

In addition, this specificity also makes it possible to envisage an increase in the extent of the bearing surfaces in the circumferential direction, and therefore to offer better retention of the blades as well as a reduction in the matting pressures.

It is noted that the assembly according to the invention is preferably designed so that the upstream and downstream bearing surfaces of one and the same blade "partially" overlap each other in the circumferential direction, in view taken according to the invention. central axis of the associated disk.

Preferably, each of the plurality of blades is designed so that in a view taken from above with respect to this blade, a principal direction in which the blade root extends from its upstream surface to its surface of downstream range, is shifted from a central axis of the disk by an angle A between 0.5 and 10 °, such as about 3 °. This then makes it possible simultaneously to obtain a satisfactory homogeneity of the intensity of the mechanical stresses encountered at the level of the bearing surfaces and stilt, and a satisfactory homogeneity of the intensity of the matting pressures encountered.

Preferably, for each of the plurality of blades, the blade root has two opposite circumferential end surfaces, arranged on either side of the bearing surfaces, these circumferential end surfaces being each substantially flat in shape . As an alternative, they can be of substantially concave shape, which makes it possible to envisage a significant increase in their extent and thus to improve the retention of the blade as well as the distribution of the matting pressures, without, however, significantly penalizing the overall mass of this dawn. Indeed, with this latter geometry, the blade root, and possibly the associated stilt, has a shape of wasp size implying that its central portion has a length in the circumferential direction less than that of the two end portions. axial axis disposed on either side of the aforementioned central portion, in the axial direction of the disc, and respectively integrating the upstream bearing surface and the downstream bearing surface.

Finally, it can be provided that each of the plurality of blades is designed so that in a view taken from above with respect to this blade, a barycentre of the upstream and downstream surfaces of the blade root, considered in this view, constitutes a center of central symmetry for upstream and downstream bearing surfaces.

The invention also relates to an aircraft engine compressor equipped with at least one such disk / blade assembly, preferably designed to form at least partially a rear stage of this compressor, and in particular a high pressure compressor .

Finally, the invention also relates to an aircraft engine, such as a turbojet, comprising at least one such compressor.

Other advantages and features of the invention will become apparent in the detailed non-limiting description below.

BRIEF DESCRIPTION OF THE DRAWINGS

• la figure 1 représente une vue en coupe d'un ensemble disque/aubes à attache marteau pour compresseur de moteur d'aéronef, selon un mode de réalisation préféré de la présente invention ;
• la figure 2 représente une vue en perspective de l'une des aubes à attache marteau faisant partie intégrante de l'ensemble montré sur la figure 1 ;
• la figure 3 représente une vue partielle de l'ensemble disque/aubes montré sur la figure 1, prise de dessus par rapport à une aube donnée de cet ensemble ; et
• la figure 4 représente une vue partielle d'un ensemble disque/aubes selon un autre mode de réalisation préféré de la présente invention, prise de dessus par rapport à une aube donnée de cet ensemble.

This description will be made with reference to the appended drawings among which;

• the figure 1 is a sectional view of a disk / hammer blade assembly for an aircraft engine compressor, in accordance with a preferred embodiment of the present invention;
• the figure 2 represents a perspective view of one of the hammer blade vanes forming an integral part of the assembly shown in FIG. figure 1 ;
• the figure 3 represents a partial view of the disk / blade assembly shown on the figure 1 taken over from a given dawn of this set; and
• the figure 4 represents a partial view of a disk / blade set according to another embodiment of preferred embodiment of the present invention, taken from above with respect to a given blade of this assembly.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

With reference first to the figure 1 it is possible to see a disk / blade assembly 1 for a high-pressure compressor of an engine of an aircraft such as a turbojet engine, this unit 1, preferably designed to form part of one of the rear stages of this high-pressure compressor. , in the form of a preferred embodiment of the present invention.

In a manner known to those skilled in the art, this set comprises first of all a disc 2 having a central axis 4 corresponding to the longitudinal axis of the turbojet. At a circumferential radial end of this disc 2, it carries a plurality of vanes 6 said hammer attachment, which are therefore distributed angularly around the central axis 4. These vanes 6 hammer clip have the specificity to include a blade root 8 intended to be housed in a circumferential groove 10 of the disk 2, the circumferential groove of the disk being located at a radial end of the disk 2 and being open radially outwardly. As is known to those skilled in the art, this circumferential groove 10 has an enlarged notch allowing the foot of each blade to penetrate into the groove, these blades being then displaced circumferentially within the groove 10. Moreover, once that all the blades have been introduced and put in place inside the groove circumferential 10, small hammers (not shown) can then be inserted to ensure the overall maintenance of all. As is clearly visible on the figure 1 , the circumferential groove 10 has generally the shape of a C opening radially outwardly, and allowing between the two ends of the C to pass the stilt of the blade as will now be described.

Indeed, each blade 6 comprises in a manner known to those skilled in the art, successively in a radial direction inward represented by the arrow 12, a blade 14, a platform 16, a stilt 18 and finally, the foot of dawn 8 supra. As such, it is noted that the blade conventionally has a leading edge 20 and a trailing edge 22, the trailing edge 22 being shifted towards the circumferential direction of the disk relative to the leading edge 20 according to a given offset direction, function of the profile of this blade. Then, the platform has a circumferential length much greater than that of the blade 14 that it supports, and is preferably intended to come closer to the platform of the two vanes 6 of the assembly that are directly adjacent thereto . Thus, when all the blades are mounted inside the groove 10, the platforms 16 of these blades substantially form a circular ring centered on the axis 4.

The stilt 18 has substantially reduced dimensions compared to those of the platform being radially outward by relative to it, both in the axial direction and in the circumferential direction of the disc. As mentioned above, this stag 18 carries radially inward the blade root 8 for the retention of the blade relative to the disk 2 on which it is mounted.

As is visible on the figures 1 and 2 , the blade root 8 can be defined as having three successive portions in the axial direction of the disk given by its central axis 4, however being noted that the entire blade root 8, and preferably the entire dawn 6, can be made in one piece, by any technique known to those skilled in the art. Thus, the blade root has indeed a central portion 26 lying generally in the internal radial extension of the stilt 18. Upstream of this central portion 26, there is an upstream axial end portion referenced 28, and having an upstream bearing surface 32 generally oriented radially outwardly. Similarly, downstream of this central portion 26, there is a downstream axial end portion referenced 30, and having a downstream bearing surface 34, also generally oriented radially outwardly.

As such, it is specified that the upstream and downstream terms used in the description are given with respect to a main flow direction of the fluid through the assembly 1, this direction being represented schematically by the arrow 40, and therefore being parallel to the axial direction of this assembly as well as to its central axis 4.

Finally, it is noted that the blade root 8 has two opposite end circumferential surfaces, respectively referenced 36, 38 on the figure 2 these surfaces being preferably located in the continuity of the opposite end-circumferential surfaces of the stilt 18, as is best seen on the figure 2 . As such, it is specified that these two surfaces 36, 38 can be substantially planar as will be described with reference to the figure 3 , and parallel to the radial direction 12 above.

As is best seen on the figure 1 it can be seen that the radial outward retention of the blade 6 with respect to the disk 2 is ensured by the contact of the two bearing surfaces 32, 34 oriented substantially radially outwards, with the two branches of the C formed by the circumferential groove 10. In this respect, it is specified that the upstream and downstream contacts sought with the bearing surfaces 32, 34 are preferably plane contacts.

Referring now to the figure 3 it is apparent from one of the features of the present invention that the upstream bearing surface 32 is offset from the downstream bearing surface 34 in the circumferential direction. More specifically, it can be seen that the trailing edge 22 of the blade 14 is shifted in the circumferential direction of the disk 2 with respect to the trailing edge 20 in a given circumferential offset direction, schematically referenced by the arrow 42 on this figure 3 . On this same figure, corresponding to a top view taken with respect to the central blade partially shown in dotted lines for the sake of clarity and located between the two blades 6 also shown in this same figure, the circumferential offset between the leading edge 20 and the edge 22 of one of these two blades located on either side of the central blade 6 has been shown schematically by the dimension referenced 44. As such, it is specifically in the same direction of circumferential shift Given that the downstream bearing surface 34 is shifted with respect to the upstream bearing surface 32, the offset here is represented diagrammatically by the dimension referenced 46.

As is clearly visible on this figure 3 , the circumferential offset of the two bearing surfaces 32, 34 is much less than that encountered between the leading edge 20 and the trailing edge 22 of the blade 14 associated. This is explained in particular by the fact that one seeks to obtain a geometry 16 according to which a main direction 48 of the blade root is offset from the central axis 4 by an angle A of between 0.5 and 10 degrees. , for example 3 degrees. It is specified that by main direction of the blade root means the direction in which this blade root extends from its upstream surface area to its downstream bearing surface, this direction may in particular be represented by a straight line by the center of gravity of each of the two abovementioned bearing surfaces, considered in view from above as shown on the figure 3 .

In this preferred embodiment of the present invention, it is effectively provided that the opposite circumferential end surfaces 36, 38 are each substantially planar in shape, ie parallel to both the radial direction of the blade and the direction. main 48 mentioned above.

As shown in figure 4 , it is possible to provide, in another preferred embodiment of the present invention, that each of these two circumferential end surfaces 36, 38 have a concave shape, thus allowing the stilt and the foot of dawn to have a global shape wasp size, allowing including an enlargement in the circumferential direction of the bearing surfaces 32, 34. In this preferred embodiment, it is expected that these concave surfaces remain substantially parallel to the direction radial of dawn. In addition, they lie in the extension of the circumferential end surfaces of the stalk 18 having the same concavity.

Whatever the preferred embodiment envisaged, it is intended to ensure that viewed from above taken from any of the blades 6, the center of gravity referenced Q on the figure 4 , corresponding to the barycentre of the upstream and downstream bearing surfaces 32, 34 combined, considered in this same view from above, constitutes a center of central symmetry for these two bearing surfaces 32, 34 associated with the same blade 6.

Of course, various modifications may be made by those skilled in the art to the invention which has just been described, only by way of non-limiting example.

Claims (8)

1. A disk/blade assembly (1) for an aircraft engine compressor, comprising a disk (2) and a plurality of blades (6) with hammer attachment mounted on said disk, each blade comprising successively, in an inward radial direction (12), an airfoil (14) comprising a leading edge (20) and a trailing edge (22) offset circumferentially from said leading edge in a given direction of offset (42), a platform (16), a stilt (18), and a blade root (8) provided with an upstream bearing surface (32) situated on a leading edge side of the airfoil and a downstream bearing surface (34) situated on a trailing edge side of this airfoil, the disk being provided with a circumferential groove (10) in which said blade root (8) of each of said plurality of blades (6) is held by means of said bearing surfaces (32, 34) resting against this circumferential groove,
   wherein, for each of said plurality of blades (6), the downstream bearing surface (34) is offset circumferentially from the upstream bearing surfaces (32) in said given direction of offset (42).
2. The di sk/blade assembly (1) for a compressor as claimed in claim 1, wherein each of said plurality of blades (6) is designed so that, in a view taken from above relative to said blade, a main direction (48) in which said blade root (8) extends, from its upstream bearing surface (32) to its downstream bearing surface (34), is offset from a central axis (4) of said disk by an angle A lying between 0.5 and 10°.
3. The disk/blade assembly (1) for a compressor as claimed in claim 2, wherein said angle A is approximately 3°.
4. The disk/blade assembly (1) for a compressor as claimed in any one of the preceding claims, wherein, for each of said plurality of blades (6), the blade root (8) has two opposite circumferential end surfaces (36, 38), arranged on either side of said bearing surfaces (32, 34), these circumferential end surfaces each having a substantially flat shape.
5. The disk/blade assembly (1) for a compressor as claimed in any one of claims 1 to 3, wherein, for each of said plurality of blades (6), the blade root (8) has two opposite circumferential end surfaces (36, 38) arranged on either side of said bearing surfaces (32, 34), these circumferential end surfaces each having a substantially concave shape.
6. The disk/blade assembly (1) for a compressor as claimed in any one of the preceding claims, wherein each of said plurality of blades (6) is designed so that, in a view taken from above relative to said blade, a baric center (Q) of said upstream and downstream bearing surfaces (32, 34) of the blade root, considered in this view, forms a central center of symmetry for said upstream and downstream bearing surfaces.
7. An aircraft engine compressor, fitted with at least one disk/blade assembly (1) as claimed in any one of the preceding claims.
8. An aircraft engine comprising at least one compressor as claimed in claim 7.

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