FR2972482A1 - Sealing sleeve for rotor of e.g. single-stage high pressure turbine of ducted-fan twin-spool turbojet of aircraft, has main body comprising strip cooperating with groove of pin of downstream flange to axially retain flange towards upstream - Google Patents

Abstract

The sleeve (102) has a main body (104) conforming to an inner surface of two directly consecutive platforms (46). The platforms are respectively placed in bases of two blades (34) of a rotor (20). The main body comprises a mechanical axial retention part i.e. strip (106), cooperating with a mechanical axial retention part i.e. groove (108), of a pin (66) of a downstream flange (60) to axially retain the flange towards upstream. The strip is radially projected from the main body towards interior or exterior of the sleeve.

Description

The present invention relates to the field of turbines for aircraft turbomachines, in particular for turbojet engines. SUMMARY OF THE INVENTION The present invention relates to the field of turbines for an aircraft turbomachine, in particular for turbojet engines. or the airplane turboprop. The invention relates more particularly to the seal between the downstream flange of a turbine stage and the blades that it contacts. STATE OF THE PRIOR ART The rotor of an aircraft turbojet turbine stage comprises a disk, vanes mounted on the disk at its periphery, and a downstream flange mounted downstream of the disk / turbine blade assembly. In known manner, the rotor is rotated by the flow of a flow of gas from upstream to downstream through the turbine. For mounting the blades, the disk is provided with peripheral teeth delimiting between them cells in which the blade roots are retained radially. At the blade platform, the turbine rotor is located close to a fixed part of the turbine, in order to close the gas flow passage, radially inward. The purpose is to prevent hot gases flowing through the turbine from escaping from the vein towards the disk. To fulfill this function, it is further provided a purge flow flowing upstream of the rotor of the turbine stage, along an upstream flange. This upstream purge flow is taken at the bottom of the combustion chamber. It is intended to follow the upstream flange before entering the vein, to prevent leakage thereof. In a conventional manner, a leakage flow from this purge flow circulates downstream between the blade platforms and the teeth of the disk. It is normally the contact between the outer radial end of the downstream flange and the blades which tends to retain this leakage flow. This restraint is crucial because it conditions the power and the proper functioning of the purge flow, since any leakage through the downstream flange impacts all the purge flow. From the prior art, it is known to reinforce the seal between the outer radial end of the downstream flange and the blades, by a particular mechanical dimensioning. More specifically, there is provided a return mass for closing the clearance in the downstream end of the flange, under the effect of the centrifugal force applied to the flange in rotation.

Although this solution makes it possible to limit the intensity of the leakage flow correctly, it nevertheless penalizes the overall mass of the turbine, because of the presence of the return mass solely dedicated to obtaining the seal at the top of the downstream flange. . In addition, this solution requires a tight tolerance on the existing game at the top of 3 downstream flange. Finally, the addition of the return mass on the downstream flange increases the level of stress within the flange, and also loads the turbine disk.

DISCLOSURE OF THE INVENTION The object of the invention is therefore to provide a simple, low-mass, economical and efficient solution to the problems mentioned above, relating to the embodiments of the prior art.

To do this, the invention firstly relates to a sealing liner for an aircraft turbine engine turbine stage rotor, said liner comprising a main body intended to fit the inner surface of two platforms respectively belonging to two directly consecutive blades of said rotor, said main body being equipped with a first mechanical member for axial retention of a flange downstream of the rotor, said first mechanical axial retention member being adapted to cooperate with a second mechanical member for axial retention of said flange downstream. The subject of the invention is also a downstream flange for an aircraft turbine engine turbine stage rotor intended to be mounted on a disk of said rotor on which vanes are mounted, said flange comprising at least one pin projecting axially towards upstream and intended to block the flange in rotation relative to the disk / blade assembly, said pin being equipped with a second axial retention member axial flange, and said second axial retention mechanical member 4 4 intended to cooperate with said first mechanical member for axial retention of the sealing sleeve. Thus, the invention cleverly provides a simple, lightweight, efficient and inexpensive mechanical solution, providing a satisfactory seal between the outer radial end of the downstream flange and the blades of the rotor. In particular, the solution adopted no longer requires a booster mass as was the case for the embodiments of the prior art. Nevertheless, a low restoring mass can be retained, without departing from the scope of the invention. The fact that the target seal is no longer obtained under the effect of the centrifugal force makes it possible in particular to obtain a reinforced seal even during the slowdown phases, during which this centrifugal force is lower. The invention applies equally well to low pressure turbines as high pressure turbines, single-stage or multistage. It applies in particular to aircraft turbojets. Preferably, said first axial retention member of the sealing sleeve takes the form of a member projecting from said main body, for example in the form of a tongue, a pin, or the like. In this case, the protruding member is preferably adapted to cooperate with a second axial retention mechanical member in the form of an orifice made in said pin, which may take a shape complementary to that of the member protrusion, such as a blind or through hole, a groove, or the like. Here, it is preferentially desired to engage the two mechanical members in order to confer the desired axial retention of the downstream flange. Preferably, the cooperation between the two mechanical members is provided to maintain the outer radial end of the downstream flange pressed against the blades of the rotor, ensuring the axial retention of the flange upstream. Preferably, said first axial retention mechanical member takes the form of a tongue protruding radially inwards or outwards, designed to fit into the second axial retention member of the flange, which then preferably takes the shape of a groove open radially outwardly or inwardly. Preferably, said first mechanical axial retention member is made in one piece with said main body. The invention also relates to an aircraft turbomachine turbine stage rotor, comprising a disk, vanes mounted on the disk, a downstream flange as described above mounted downstream of the disk and vanes, and at least one sealing liner as mentioned above. The invention also relates to an aircraft turbine engine turbine comprising at least one turbine stage rotor as described above. Finally, the subject of the invention is a turbomachine for an aircraft, comprising at least one such turbine. Other advantages and features of the invention will become apparent in the detailed non-limiting description below. BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood, and other details, advantages and characteristics thereof will appear on reading the following description given by way of non-limiting example and with reference to the appended drawings in which: Figure 1 is a partial schematic half-view in axial section of an aircraft turbojet comprising a high pressure turbine according to a preferred embodiment of the present invention; FIG. 2 is a partial perspective view of the rotor of a stage of the high pressure turbine shown in the preceding figure, the sealing jacket having been removed for reasons of clarity; FIG. 3 is a partial schematic half-view in axial section of the rotor shown in the previous figure; FIG. 4 is a perspective view of one of the rotor blades of the rotor shown in FIGS. 2 and 3, shown exploded with its associated sealing jacket; FIG. 5 represents a view similar to that of FIG. 2, showing more precisely the second axial retaining member provided on one of the anti-rotation pins of the downstream flange; and - Figure 6 shows a partial view similar to that of Figure 3, according to an alternative embodiment. In these figures, identical or similar elements are designated by identical reference numerals. DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT Referring firstly to FIG. 1, it is possible to see an aircraft turbojet 1 of the double-flow and double-body type. The turbojet comprises, successively in the direction of thrust represented by the arrow 2, a low pressure compressor 4, a high pressure compressor 6, a combustion chamber 8, a high pressure turbine 10 and a low pressure turbine 12. The high pressure turbine 10, of the single-stage type, is in the form of a preferred embodiment of the present invention, which will now be detailed with reference to FIGS. 2 to 5. The turbine comprises a stage rotor 20 which located downstream of a high pressure distributor 24, belonging to the previous stage. Downstream of the rotor 20 is a fixed portion 22 of the turbojet, corresponding to the inlet of the low pressure turbine. First, the rotor 20 has a disk 26 whose radially outer periphery has teeth 28 circumferentially spaced from each other. Blocks 30 are defined 8 between the teeth of the disc. They are axial or oblique, open on both the upstream face 31 and the downstream face 33 of the disk 26. In a conventional manner, each cell 30 receives the root 32 of a blade 34, in order to retain it radially towards the outside, by form cooperation. The root 32 of each vane 34 is located at a distance from the bottom 36 of its associated cell 30, in the radial direction. Thus, there is formed a free space 38 between the end of the foot 32 and the bottom of the cell 36. At the upstream end, the foot 32 may possibly extend radially inwards by a first wall 40, as well as downstream, the foot 32 extends radially inwards by a second wall 42. In addition, each blade 34 comprises at its base a platform 46 intended in particular to guide the flow from upstream to downstream in the turbine, that is to say from the inlet to the outlet of the turbine, a primary flow of gas from the combustion chamber. This platform 46, arranged between the foot 32 and the aerodynamic blade 48 of the blade, extends circumferentially on either side of this blade 48. It is classically in the extension of an identical platform belonging to a blade. directly consecutive, as is visible in Figure 2. The rotor 20 includes an upstream flange 44 30 carried by the disc 26 at the upstream face 31 thereof. The annular flange 44 has a periphery equipped with hooks 50 spaced circumferentially from each other, each hook cooperating with an external radial projection 52 provided in front of each disc tooth 28. Circumferential notches 53 therefore appear between the hooks 50 , axially facing the blade roots, as is best seen in FIG. 3. In this respect, it is indicated that each blade 34 incorporates an upstream support tongue 55 of the platform 46, which extends under the platform according to a circumferential length greater than that of the foot bearing this tongue. However, its circumferential length is smaller than that of the platform, in order to define between the tabs 55 of the consecutive blades, circumferential slots 57 traversed by the hooks 50. On the inner radial portion, the flange 44 has an annular hook 54 cooperating with a complementary hook 56 made on the upstream face 31 of the disc. The upstream flange 44 in particular participates in the axial retention of the blades in the cells, towards the front. The rotor 20 also incorporates a downstream flange 60 carried by the disk 26 at the downstream face 33 thereof. The annular flange 60 has a periphery intended to be in abutment against a tongue 62 of downstream support of the platform 46, which extends under the platform along a circumferential length greater than that of the foot which bears this tongue 62, this length being on the other hand substantially identical to that of the platform. Free spaces 64 are then created located under the tongues 62, delimited circumferentially by the upper part of the blade roots, and delimited radially between the tongues 62 and the upper face of the disk teeth 28. The spaces 64 are traversed axially by pins 66 protruding axially upstream from an upstream face 68 of the downstream flange. This makes it possible to confer an anti-rotation function on the downstream flange with respect to the disk / blade assembly. As an indication, the number of pins 66 is much smaller than the number of spaces 64, for example only one space 64 of twenty being traversed by such a pin 66.

The pins 66 are located near the periphery of the flange, radially inwardly relative to the outer radial end 80 of the same downstream flange 60. A seal 71 carried by the upstream face 68 of the flange 60 contact the downstream face 33 of the disk 26. The internal radial delimitation is obtained by contact between the upstream face 68 of the flange 60 and the downstream face 33 of the disk 26, under the spaces 38.

In operation, a purge flow 76 flows along the upstream flange 44, radially outwardly. This purge flow 76, taken from the bottom of the combustion chamber, is intended to be introduced into the vein between the fixed combustion chamber outlet portion 24 and the upstream end of the blade platforms, in order to avoid the leaks from the vein. From this purge flow 76 is output a leakage flow 78 flowing downstream under the blade platforms, between these and the teeth of the disc. To enter the space under the platforms 46, the air passes through the notches 53 of the upstream flange 44 and the circumferential slots 57 of the blades, above the hooks 50. To best limit this leakage flow 78 which tends to escape between the outer radial end 80 of the downstream flange 60 and platform supports 62 of the blades 34, it is desired to have a satisfactory contact intensity between these elements 62, 80. To do this, the flange downstream 60 is retained axially, upstream, that is to say, prevented from moving axially downstream, with the aid of a sealing liner 102 associated with the two vanes 34 between which penetrates the counter 66 of the flange 60. Indeed, the sealing sleeve 102 comprises a conventional main body 104 conforming to the inner surface of the two platforms 46 directly consecutive. It takes a rectangular overall shape whose periphery is curved radially inwards, as can be seen in FIGS. 3 and 4. At its small downstream side, the main body 104 carries a first mechanical member 106 for axial retention of the flange 60 , which here preferably takes the form of a tongue projecting radially inwards from the small convex downstream side of the central body 104. The tongue 106, preferably made in one piece with the body 104 , extends over a circumferential length less than or equal to that of the pin 66 with which it cooperates directly.

Indeed, the latter is equipped with a second mechanical member 108 for axial retention of the flange 60, which here preferably takes the form of a groove open radially outward, and receiving the tongue 106. The groove 108, which is practiced on the radially outer face of the pin 66, also extends over a circumferential length less than or equal to that of the pin, and preferably the same or similar to that of the tab 106 that it receives by interlocking.

The assembly is carried out first by assembling the upstream flange on the disc, then mounting all the blades equipped with sealing sleeves on the same disc. Finally, the downstream flange is mounted on the disk, so as to create an interlocking tabs 6 in the corresponding grooves 108. Thus, the sealing liner 102, the main function of which is to limit leakage in the inter-platform clearance that it covers, fulfills an additional function of axial retention upstream of the flange 60. This specificity can be attached. to one, several, or all of the pins 66 and their corresponding protective shirts. According to an alternative embodiment 30 shown in FIG. 6, the first mechanical member 106 takes the form of a projecting tongue 13 radially outwardly. It constitutes one of the two branches of a U integral with the main body of the liner 104, with which this U is preferably made in one piece. The U is penetrated by the tongue 62 of downstream support of the platform 46, which is arranged between the two branches of the U, radially opposite the bottom of this U through the free space 64. The second mechanical member axial retention is here also a groove 108, open radially inwards, and defined by the outer radial end 80 of the downstream flange 60 in the form of a hook. The interlocking of the tongue 106 in the groove 108, by centrifugal effect, makes it possible to directly press the end 80 in the form of a hook against the downstream face of the platform support tongue 62. This interlocking advantageously takes place at a low speed. due to the low mass of the tongue 106 subjected to the centrifugal force.

In this embodiment of FIG. 6, the anti-rotation function of the downstream flange 60 is also conferred by a plurality of pins identical or similar to the pins 66 of the embodiment of FIG. 3, for example three pins (not shown on FIG. FIG. 6) respectively inserted in three associated spaces 64. The pins can then carry a groove housing a tab secured to the main body of the liner 104, as was the case for the previous embodiment. Thus, the few spaces 64 concerned 14 are only traversed by the anti-rotation pins, but not by the aforementioned U. 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 sealing jacket (102) for an aircraft turbine engine turbine stage (10) rotor (20), said jacket comprising a main body (104) for conforming to the inner surface of two platforms (46) respectively belonging to two blades (34) directly consecutive of said rotor, characterized in that said main body (104) is equipped with a first mechanical member (106) for axial retention of a downstream flange (60) according to any one of claims 5 at 7, said first mechanical axial retention member (106) being intended to cooperate with the second mechanical axial retention member (108) belonging to said downstream flange. 2. Sealing sleeve according to claim 1, characterized in that said first axial retention member (106) takes the form of a member protruding from said main body (104). Sealing sleeve according to claim 2, characterized in that said first axial retention member (106) takes the form of a radially protruding tongue, inwardly or outwardly. 4. Sealing sleeve according to any one of the preceding claims, characterized in that said first mechanical retention member 16 axial (106) is formed integrally with said main body (104). 5. Downstream flange (60) for rotor (20) turbine stage (10) of an aircraft turbine engine, intended to be mounted on a disk (26) of said rotor on which vanes (34) are mounted, said flange comprising at least one pin (66) protruding axially upstream and intended to block this flange in rotation relative to the disk / blade assembly, characterized in that said pin (66) is equipped with a second mechanical member Axial retaining member (60), said second axial retention member (108) being adapted to cooperate with said first axial retention member (106) belonging to the sealing sleeve (102) according to any of claims 1 to 4. 6. Downstream flange according to claim 5, characterized in that said second axial retention mechanical member (108) takes the form of an orifice formed in said pin. 7. Downstream flange according to claim 6, characterized in that said second axial retention member (108) takes the form of a radially open groove, outwardly or inwardly. An aircraft turbine engine turbine stage (10) rotor, comprising a disk (26), vanes (34) mounted on the disk, a downstream flange (60), any one of claims 5, to 7 mounted downstream of the disk and blades, and at least one sealing sleeve 17 (102) according to any one of claims 1 to 4. An aircraft turbine engine turbine (10) comprising at least one turbine stage rotor (20) according to claim 8. 10. Turbine engine (1) for aircraft, characterized in that it comprises at least one turbine (10) according to claim 9.15