AIRCRAFT TURBOMACHINE ASSEMBLY INCLUDING A COVER RING FOR ISOLATING MECHANICAL FIXINGS FROM AIRFLOW

Abstract

The invention relates to an assembly (50) for an aircraft turbomachine comprising a first element (25) and a second element (35), each centered on a central longitudinal axis (2). The first element (25) partially defines an air circulation cavity (46). The assembly also includes a mechanical linkage (42) between the first and second elements (25, 35). This mechanical linkage comprises a plurality of mechanical fastening members (36) distributed around the central longitudinal axis (2). An end portion (38) of each of these mechanical fastening members (36) is axially projecting from the first element (25). According to the invention, the assembly further comprises, fixed to the first element (25), a ring (52) for covering the end portions (38) of the mechanical fastening members (36). The ring (52) defines a sub-cavity (46a) for isolating the end portions (38). Figure for the abbreviation: Figure 4.

Description

Title of the invention: AIRCRAFT TURBOMACHINE ASSEMBLY COMPRISING A COVER RING FOR ISOLATING MECHANICAL FIXINGS FROM AIRFLOW Technical field

[0001] The invention relates to the field of aircraft turbomachinery, and more particularly to assemblies comprising mechanical fastening elements, such as bolts, the end parts of which penetrate into air circulation cavities. Such bolts are disclosed, for example, in document EP 1 637 701 A1.

[0002] The invention applies to any type of aircraft turbomachine, such as turbojets and turboprops. STATE OF PRIOR ART

[0003] Aircraft turbomachinery usually comprises several assemblies of two elements fastened to each other by means of mechanical fasteners, the end portions of which penetrate an airflow cavity delimited by one of these two elements. The mechanical fasteners, such as bolts, are generally distributed around the longitudinal central axis, forming an annular row of fasteners.

[0004] The presence of the end portions of the fastening elements within the airflow cavity generates a phenomenon of air rotation around each of these protruding end portions. This phenomenon occurs because the annular row of fastening elements, taken as a whole, does not constitute an axisymmetric element. The rotation of the air, locally at the level of each mechanical fastening element, consumes energy and thus reduces the turbine's efficiency, thereby increasing engine fuel consumption. Depending on the application, the end portions of the fastening elements can also slow down / disturb the already rotating air, with the same detrimental consequences on engine efficiency and fuel consumption.

[0005] By way of example, the assembly may include a turbine or compressor wheel disc, connected by bolts to an inter-disc connecting ferrule. In this case, it is the screw heads or nuts that protrude into the air circulation cavity delimited by the disc, and which generate the detrimental phenomenon described below, also known as "churn".

[0006] There therefore remains a need to improve the design of these together, in order to reduce the negative impacts on specific fuel consumption. Summary of the invention

[0007] To meet at least partially this need, the invention relates to an assembly for an aircraft turbomachine comprising a first element and a second element each centered on a central longitudinal axis, the first element partially delimiting an air circulation cavity and the assembly also comprising a mechanical link between the first and second elements, the mechanical link comprising a plurality of mechanical fixing members distributed around the central longitudinal axis, an end part of each of these mechanical fixing members being axially projecting relative to the first element.

[0008] According to the invention, the assembly further comprises, fixed on the first element, a ring covering the end parts of the mechanical fastening members, the ring delimiting with the first element a sub-cavity for isolating the end parts of the mechanical fastening members from the air intended to circulate in said cavity.

[0009] In other words, the end portions of the mechanical fastening elements are advantageously confined within the sub-cavity delimited by the cover ring, so that the air circulating in the cavity delimited by the first element is no longer disturbed by these end portions, or at least less so than in prior art designs. In particular, the churning phenomenon is greatly reduced, or even completely eliminated, resulting in a gain in specific fuel consumption.

[0010] The invention preferably provides for at least one of the following optional technical features, taken individually or in combination.

[0011] Preferably, the ring forms a cover over the mechanical fastening elements, the cover being free of discontinuity around the longitudinal central axis (therefore in a circumferential direction of the assembly), so as to limit as much as possible any disturbances on the airflow circulating in the cavity.

[0012] Preferably, the assembly includes means for mounting the cover ring onto the first element, these mounting means being free of screw-on elements. Preferably, these mounting means for the cover ring include a dog clutch connection, or an axial retaining ring, or any other technical solution for limiting protrusions in the air circulation cavity.

[0013] Preferably, the mechanical fasteners are bolts.

[0014] Preferably, the first element is a turbine wheel disc or a com press, and even more preferably a rotating wheel disc of a turbine or compressor, subjected to a high-pressure primary flow. Indeed, since the churning phenomenon is all the more significant as the rotational speed of the wheels is high, it proves particularly advantageous to implement the invention on high-pressure components, which rotate at the highest speeds and on which the expected gains are therefore the greatest.

[0015] Preferably, the assembly comprises an annular row of blades mounted on the periphery of the disk in grooves for housing these blades, the cover ring comprising a peripheral portion forming axial retention of the blades relative to the disk, and the peripheral portion of the ring defining a sub-cavity for supplying air to the blade housing grooves. In this preferred embodiment, the cover ring thus advantageously fulfills one or more additional functions.

[0016] Preferably, the second element is an inter-disc connecting ferrule.

[0017] Preferably, the ring axially covers the end portions of the mechanical fastening members, as well as radially on either side of these end portions. Preferably, the ring comprises a portion axially opposite and at a distance from the end portions, from which two branches extend radially opposite the end portions, the branches being preferably in watertight contact with the first element.

[0018] The invention also relates to an aircraft turbomachine turbine or compressor comprising at least one assembly as described above, and preferably a high-pressure turbine or compressor.

[0019] Finally, the invention also relates to an aircraft turbomachine comprising at least one assembly as described above, the turbomachine preferably being a twin-body and / or twin-body turbojet.

[0020] Other advantages and features of the invention will appear in the detailed, non-limiting description below. Brief description of the drawings

[0021] This description will be made with reference to the attached drawings, among which;

[0022] [Fig.1] is a schematic axial cross-sectional view of a turbofan engine according to the invention;

[0023] [Fig.2] represents a half-axial cross-sectional view of a part of the high-pressure turbine of the turbojet shown in the previous figure, this part of the turbine comprising an assembly according to a first preferred embodiment of the invention;

[0024] [Fig.3] represents a perspective view of a moving wheel of the high-pressure turbine shown in the previous figure;

[0025] [Fig.4] is a perspective view similar to the previous one and showing the whole of turbine, cut along a longitudinal plane;

[0026] [Fig.5] is a view similar to that of [Fig.3], with the cover ring shown partially and exploded;

[0027] [Fig. 6] represents a partial axial cross-sectional view, showing an alternative embodiment for the means of mounting the cover ring on the turbine wheel disc; and

[0028] [Fig.7] represents a view similar to that of [Fig.4], representing the whole according to a second preferred embodiment of the invention. DETAILED EXPLANATION OF PREFERRED METHODS OF IMPLEMENTATION

[0029] With reference first to [Fig. 1], an aircraft turbomachine 1 is shown, according to a preferred embodiment of the invention. This is a twin-spool, turbofan engine. However, it could be a turbomachine of another type, for example a turboprop, without departing from the scope of the invention.

[0030] The turbomachine 1 has a central longitudinal axis 2 around which its various components extend. It comprises, from upstream to downstream along a main direction 5 of gas flow through this turbomachine, a blower 3, a low-pressure compressor 4, a high-pressure compressor 6, a combustion chamber 11, a high-pressure turbine 7 and a low-pressure turbine 8.

[0031] Conventionally, after passing through the blower, the air splits into a central primary flow 12a and a secondary flow 12b which surrounds the primary flow. The primary flow 12a flows into a main gas circulation channel 14a passing through the compressors 4, 6, the combustion chamber 11 and the turbines 7, 8. The secondary flow 12b, on the other hand, flows into a secondary channel 14b delimited radially outwards by an engine casing, surrounded by a nacelle 9.

[0032] Fig. 2 represents a part of the high-pressure turbine 7, on which is shown successively, from upstream to downstream along the direction 5, an upstream turbine wheel 20, a fixed wheel 22 forming a distributor, and a downstream wheel 24. Preferably, these are the first stages of the high-pressure turbine 7.

[0033] Each movable wheel 20, 24 comprises a disc 25 centered on the axis 2, at the periphery of which is an annular row of blades 26, the feet of which 28 are housed in blade housing grooves 30 made through a peripheral part 32 of the disc.

[0034] The two discs 25 are connected to each other by an inter-disc connecting ferrule 34, extending axially inside the fixed wheel 22. At the upstream end of the inter-disc connecting ferrule 34, it has a mounting flange 35 centered on the axis 2 and through which bolts 36 pass, these also passing through a radially external portion of the disc 25 of the upstream moving wheel. turbine 20. At its downstream end, the ferrule 34 is preferably made in one piece with the disc 25 of the downstream rotating wheel 24.

[0035] In the example shown in [Fig.2], the screw head 38 of each bolt 36 is axially supported against the disc 25 of the movable wheel 20, while the nut 40 is axially supported against the fixing flange 35. However, an inverse configuration could be adopted, without departing from the scope of the invention.

[0036] The bolts 36, arranged in a ring centered on the axis 2, thus form a mechanical connection 42 between the disc 25 and the flange 35, with the screw heads 38 forming upstream end portions projecting axially upstream relative to this disc 25 of the turbine wheel 20. As is known, the disc 25 of the turbine wheel 20 partially delimits, axially downstream, an air circulation cavity 46. The airflow has been schematically represented in [Fig. 2] by the arrow 48. It conventionally originates from an air intake on one of the compressors 4, 6 or on the secondary flow 14b, then passes through the cavity 46 before heading towards the bottom of the grooves 30 of the disc, so as to cool the blades 26 housed in these grooves.

[0037] The disc 25 of the upstream moving wheel of the turbine 20, the interdisc connecting ferrule 34, and the bolts 36 form an assembly 50 specific to the present invention, a first preferred embodiment of which will be described jointly with reference to figures 2 to 5.

[0038] First, it is noted that direction 5 also corresponds to the main direction of airflow within assembly 50. According to a particular feature of the invention, the latter comprises a cover ring 52 for the screw heads 38, fixed to the disc 25 of the wheel 20, by being located upstream of this disc 25. The cover ring 52, centered on the axis 2, delimits with the disc 25 a sub-cavity 46a for isolating the screw heads 38 from the air circulating in the cavity 46. In other words, instead of being exposed to the airflow 48 circulating in the cavity 46 and potentially generating harmful churning phenomena, the screw heads 38, projecting upstream, are isolated in the closed annular sub-cavity 46a. in order not to disrupt this airflow.

[0039] Also to limit disturbances in the airflow through the cavity 46, the cover ring 52 is essentially made up of a hood 56 free from discontinuity around the axis 2, that is to say in the circumferential direction 58 of the assembly 50, this hood 56 being intended for covering the mechanical fastening elements 36.

[0040] This annular-shaped cover 56 thus constitutes the greater part of the ring 52, being located axially opposite the screw heads 38. At its outer radial end, the cover 56 has a curved portion forming an axial arm 60, bearing radially against a recess in the peripheral part 32 of the disc 25, so as to close the sub-cavity 46a radially outwards. The junction between the peripheral part 32 of the disc and the cover ring 52 is located radially inwards relative to the bottom of the grooves 30 housing the blades, so that this ring 52 does not impede the evacuation of the airflow 48 downstream through these same grooves 30.

[0041] In addition, at its radially internal end, the annular cover 56 has another branch 60' which is integral with means for mounting the cover ring 52 on the disc 25.

[0042] Thus, the hood 56 has a radial portion 59 axially opposite the screw heads 38, and at the ends of which extend the two branches 60, 60', each radially opposite the heads 38. In addition, the branches 60, 60' are in tight contact with the disc 25, to ensure the confinement of the screw heads 38 in the dedicated sub-cavity 46a.

[0043] The ring mounting means, preferably free of screw-on elements such as bolts, comprise a dog clutch 62 formed with teeth 64 projecting radially inward from the radially inner end of the cover 56. The clutch 62 is completed by a radial collar 66 integral with the disc 25, provided with notches 68 opening radially outward and sized to allow the axial insertion of the teeth 64. The number of notches 68, preferably the same as the number of teeth 64, may preferably be between four and ten. This dog clutch 62, which closes the sub-cavity 46a radially inward, causes only a negligible disturbance to the airflow 48 passing through the cavity 46, or even no disturbance at all.

[0044] Consequently, the cover 56 covers axially the screw heads 38, as well as radially on either side of these heads.

[0045] Figure 6 shows an alternative embodiment for the mounting means of the cover ring 52 on the disc 25. In this alternative, the dog clutch connection is replaced by a connection comprising an axial retaining ring 70, preferably split. The retaining ring 70 is housed in a groove 72 defined between the radial flange 66 and the flank of the disc 25, and bears axially downstream on a bearing surface 74 of the radially internal end of the cover 56 of the cover ring 52.

[0046] In a second embodiment shown in [Fig. 7], the cover 56 of the cover ring 52 is extended radially outwards, so as to include a peripheral portion 76 forming axial retention of the blade feet 28 in the grooves 30 of the disc 25. Furthermore, to further functionalize this ring 52, its peripheral portion 76 can define a sub-cavity 46b for supplying air to the blade housing grooves 30, since the sub-cavity 46b opens onto the bottom of the blade housing grooves 30.

[0047] The sub-cavity 46b, preferably annular and centered on the axis 2, is itself supplied with air from the cavity 46, through axial openings 80 made through the peripheral part 76.

[0048] Of course, various modifications can be made by a person skilled in the art to the invention just described, solely by way of non-limiting examples, and the scope of which is defined by the attached claims.

Claims

Claims

1. Assembly (50) for an aircraft turbomachine comprising a first element (25) and a second element (35) each centered on a longitudinal central axis (2), the first element (25) partially delimiting an air circulation cavity (46) and the assembly also comprising a mechanical connection (42) between the first and second elements (25, 35), the mechanical connection comprising a plurality of mechanical fixing members (36) distributed around the longitudinal central axis (2), an end portion (38) of each of these mechanical fixing members (36) projecting axially relative to the first element (25), the first element (25) being a turbine or compressor wheel disc, and the second element being an inter-disc connection ferrule (35), characterized in that it further comprises, fixed to the first element (25), a ring (52) for covering the end portions (38) of the mechanical fixing members (36),the ring (52) delimiting with the first element (25) a sub-cavity (46a) for isolating the end parts (38) of the mechanical fixing members (36), with respect to the air intended to circulate in said cavity (46).,

2. Assembly according to claim 1, characterized in that the ring (52) forms a cover (56) covering the mechanical fixing members (36), the cover (52) being free from discontinuity around the longitudinal central axis (X).

3. Assembly according to claim 1 or 2, characterized in that it comprises mounting means (62, 70) of the covering ring (52) on the first element (25), these mounting means being devoid of screwed elements.

4. Assembly according to claim 3, characterized in that the means for mounting the covering ring (52) comprise a dog connection (62), or an axial retaining ring (70).

5. Assembly according to any one of the preceding claims, characterized in that the mechanical fixing members are bolts (36).

6. Assembly according to any one of the preceding claims, characterized in that it comprises an annular row of blades (26) mounted at the periphery of the disc in housing grooves (30) for these blades, the covering ring (52) comprising a peri- spherical (76) forming axial retention of the blades (26) relative to the disc (25), and in that the peripheral part (76) of the ring defines a sub-cavity (46b) for supplying air to the blade housing grooves (30).

7. Assembly according to any one of the preceding claims, characterized in that the ring (52) axially covers the end portions (38) of the mechanical fixing members (36), and in that the ring (52) further covers radially on either side the end portions (38), the ring (52) preferably comprising a portion (59) axially facing and at a distance from the end portions (38), portion (59) from which two branches (60, 60') extend radially facing the end portions (38), the branches (60, 60') being in contact with the first element (25).

8. Turbine (7, 8) or compressor (4, 6) of an aircraft turbomachine comprising at least one assembly (50) according to any one of the preceding claims.

9. Aircraft turbomachine (1) comprising at least one assembly (50) according to any one of claims 1 to 7.

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