Disc for a bladed rotating wheel of an aircraft turbomachine module, made using several disc parts in different materials

Abstract

The invention relates to a disc (20) for a bladed rotating wheel (12B) of an aircraft turbomachine module, the disc extending circumferentially around a central longitudinal axis (A1), and comprising: - a first part (22a) made of a first material, and having teeth (26) between which are defined recesses for housing the blades (30) of the rotating wheel; - a second part (22b) made of a second material different from the first material, the second part being assembled to the first part (22a), and defining at least partially a disc bore (44) centered on the central longitudinal axis (A1). Figure for the abstract: Fig. 2

Description

Title of the invention: DISC FOR A SPLINED MOVABLE WHEEL OF AN AIRCRAFT TURBOMACHINE MODULE, MADE USING SEVERAL DISC PARTS MADE OF DIFFERENT MATERIALS Technical field

[0001] The present invention relates to an aircraft turbomachine, preferably of the turbojet or turboprop type.

[0002] More particularly, the invention relates to the design of a disc for a bladed rotating wheel of an aircraft turbomachine module, for example a turbine. This type of disc is known for example from document FR 2 825 748 A1. Prior art

[0003] A disc of a bladed rotating wheel of an aircraft turbomachine module is generally made in one piece, usually from a raw casting.

[0004] The disc is a fairly large part with relatively complex shapes, notably to create recesses for housing the blade roots, or a central bore for the passage of the drive shaft(s). The as-cast part must meet certain quality requirements to be retained for the subsequent manufacturing process, which includes machining. For example, if there are defects in the material of the as-cast part, it may have to be rejected entirely. The larger the part, the greater the risk of such defects.

[0005] Furthermore, the shape and dimensions of the disc are complicated by the fact that it has different zones subjected to various stresses. For example, on the outer periphery of the disc, the teeth forming the grooves are located close to the turbomachine's gas stream and are thus potentially subjected to high temperatures. This stress locally exposes the disc to a risk of wear. Another example is found on the inner periphery of the disc, which defines the bore for the drive shafts. This bore constitutes a zone of accelerated aging, as it exhibits lower fatigue strength.

[0006] Every material is more or less adapted to each of these local constraints on the disk. This leads, during the design phase, to the need to plan for specific shapes and / or dimensions, taking into account the material chosen for the disk. The resulting design therefore still needs improvement. Description of the invention

[0007] To address at least partially the drawbacks mentioned above relating to prior art inventions, the invention first relates to a disc for a bladed rotating wheel of an aircraft turbomachine module, the disc extending circumferentially around a longitudinal central axis, and comprising:

[0008] - a first part made of a first material, and having teeth between which are defined cells for housing the blades of the moving wheel;

[0009] - a second part made of a second material different from the first material, the second part being assembled to the first part, and defining at least partially a central disc bore centered on the longitudinal central axis.

[0010] The invention advantageously provides for manufacturing the disc in several parts formed from different materials. Each part of the disc can thus be made from a blank, for example, a casting, of smaller dimensions than in the prior art. This reduces the risk of defects in the material, and if disposal is necessary, the volume of the defective blank is advantageously smaller, thereby resulting in limited loss.

[0011] Furthermore, dissociating the disc into several assembled parts allows for optimal adaptation of the material for each part, according to local constraints. For example, for one part, it may be advantageous to prioritize abrasion resistance over fatigue resistance, and vice versa for another part of the disc, resulting in an increased lifespan for the entire disc. The disc design can thus be improved in terms of shape and dimensions, leading to a reduction in its overall mass. As such, the invention therefore represents a result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of these aircraft (decarbonization).

[0012] The invention preferably provides for at least one of the following optional technical features, implemented individually or in combination.

[0013] Preferably, the first and second parts are annular, centered on the central longitudinal axis. Alternatively, these parts could constitute only annular sectors of the disk, without departing from the scope of the invention.

[0014] Preferably, the disc also includes a third part, preferably made of the second material, and also defining at least partially the disc bore.

[0015] Preferably, the second and third parts axially enclose the first part of the disk, the first part comprising a radially internal portion of fixing, axially clamped between two radially external fixing portions belonging respectively to the second and third parts, the radially internal and external fixing portions being equipped with radial retention means of the first part relative to the second and third parts.

[0016] Preferably, the second and third parts each have the shape of a disc with a hub and a rim, and they preferably constitute two symmetrical or substantially symmetrical parts with respect to a radial mounting interface plane between these two parts.

[0017] Preferably, the disc includes means for fixing the second part to the third part, said fixing means also preferably passing through the first part, in order to reinforce the assembly between the parts of the disc.

[0018] Preferably, as mentioned above, the first material has superior strength to fining than the second material, and / or the second material has superior strength to fatigue than the first material.

[0019] Preferably, the radial retention means of the first part, relative to the second and third parts, comprise on the one hand annular ridges or ribs, and, on the other hand, annular grooves configured to cooperate in a complementary manner with the annular ridges or ribs.

[0020] Preferably, the radial retaining means are arranged radially on the inside and radially on the outside with respect to said fastening means.

[0021] The invention also relates to a bladed movable wheel of an aircraft turbomachine module, comprising such a disc, as well as movable blades housed in the recesses of the first part, the movable wheel also comprising, preferably, a retaining flange for an axial retaining ring of the movable blades.

[0022] Preferably, the means for fixing the second part to the third part also pass through the retaining flange of the axial retention ring, in order to make the wheel even lighter and more compact.

[0023] The invention also relates to an aircraft turbomachine module comprising at least one such rotating wheel, the module preferably being a turbine, and even more preferably a low-pressure turbine. The application of the invention to a compressor also remains possible.

[0024] Finally, the invention also relates to an aircraft turbomachine, comprising at least one such module.

[0025] Other advantages and features of the invention will become apparent from the following detailed, non-limiting description. Brief description of the drawings

[0026] The detailed description that follows refers to the accompanying drawings in which:

[0027] [Fig-1] is a schematic axial cross-sectional view of a turbojet engine according the invention;

[0028] [Fig.2] is a half-axial cross-sectional view of part of a turbine stage of the turbomachine shown in the previous figure, the turbine stage comprising a moving wheel being in the form of a preferred embodiment of the invention;

[0029] [Fig.3] is an enlarged perspective view of a first part of the disk shown in [Fig.2];

[0030] [Fig.4] is an enlarged perspective view of a second and a third part of the disk shown in [Fig.2]; and

[0031] [Fig. 5] is an enlarged perspective view of the disk obtained by assembling the first, second, and third disk parts shown in Figures 3 and 4. Detailed description of embodiments

[0032] The figures include a reference frame L, R and C defining respectively longitudinal, radial and circumferential directions orthogonal to each other, these directions corresponding to those of an aircraft turbojet 1 according to the invention.

[0033] Figure 1 represents the aircraft turbojet engine 1, preferably having a twin-spool, twin-flow design. However, other types of turbojet engines are possible, as are turbomachinery other than turbojet engines, such as turboprop engines.

[0034] Hereafter, the terms "upstream" and "downstream" are defined with respect to a principal direction DI of gas flow through the turbojet 1 when it is operating in direct thrust mode. The direction DI is parallel to the longitudinal direction L, and also parallel to a longitudinal axis Al of the turbojet, around which its various components extend. In this case, from upstream to downstream of the turbojet 1, these components are a fan 4, a low-pressure compressor 5, a high-pressure compressor 6, a combustion chamber 7, a high-pressure turbine 8, and a low-pressure turbine 9.

[0035] During the operation of the turbojet 1, an airflow 10 enters the turbojet 1 through an air inlet 3, passes through the fan 4, and then splits into a central primary flow 10A and a secondary flow 10B. The primary flow 10A flows in a main gas circulation channel 11A passing through the compressors 5 and 6, the combustion chamber 7, and the turbines 8 and 9. The secondary flow 10B, on the other hand, flows in a secondary channel 11B surrounding the main channel 11A, also called the primary channel, or aerodynamic flow channel.

[0036] In a manner known per se, a turbine such as the high-pressure turbine 8 or the low-pressure turbine 9 comprises one or more stages 12. Each stage 12 includes a stator assembly 12A comprising a distributor 29, also called a bladed stator ring, and a moving wheel 12B, arranged directly upstream of the assembly 12A. Such a stage 12 is shown in [Fig.2].

[0037] With reference now to figures 2 to 5, the turbine wheel 12B will be described, here preferably integrated into the high-pressure turbine.

[0038] One of the features of the invention lies in the fact that the rotating wheel 12B comprises a turbine disc 20 centered on the axis Al, around which it extends in the direction C, and made by assembling several parts of different materials. In the preferred embodiment described below, these are three separate parts assembled together, each obtained separately, for example from as-cast pieces.

[0039] More specifically, the disc 20 first comprises a first portion 22a, annular in shape, centered on the axis A1, and made of a first material. This first portion 22a forms the outer periphery of the disc 20, defining in particular radial teeth 26 between which are defined recesses 28 for housing blades 30 of the rotating wheel. The recesses 28 are thus circumferentially spaced from one another, each being axially open, preferably in both directions. In a known manner, each recess 28 adopts a shape adapted so as to be able to radially house and retain the foot of a blade 30, of complementary shape. To this end, the blade foot is, for example, bulbous, or it forms with the recess 28 a so-called "fir tree" attachment.

[0040] The first portion 22a comprises a support 32 for the teeth 26, the support 32 being annular and located radially inward with respect to the teeth 26. The support 32 thus forms the radial base of the sockets. In addition, the first portion 22a comprises, projecting radially inward from the tooth support 32, a radially internal attachment portion 34. This attachment portion 34 has an axial thickness less than that of the support 32. The axial thickness is defined between the two opposing surfaces 36 of the attachment portion 34, each equipped with radial retention elements 40 for this first portion 22a of the disc. The retention elements 40 are here annular ridges or ribs, or any other form of axial projection.

[0041] In addition, the radially internal fixing portion 34 is traversed by through holes 42, axially oriented and circumferentially spaced from each other.

[0042] The disc 20 also comprises two other parts, namely a second part 22b and a third part 22c. These two parts 22b, 22c are preferably symmetrical or substantially symmetrical in shape with respect to a median transverse plane of the disc and the wheel, which also corresponds to a radial interface plane between these two parts. Furthermore, the second and third parts 22b, 22c each have the shape of a disc with a hub and a rim, and they constitute preference two parts symmetrical or substantially symmetrical with respect to a radial mounting interface plane between these two parts 22b, 22c.

[0043] Therefore, in what follows, only one of the second and third parts 22b, 22c will be described, while in the figures, the symmetrical portions of the two parts 22b, 22c will bear numerical references with the extension "b" for the second part 22b, and with the extension "c" for the third part 22c.

[0044] Thus, each of the second and third parts 22b, 22c has an annular shape centered on the axis Al, and is made of a second material different from the first material used for the first part. Each of them defines a part 44b, 44c of a central bore of the disk 44 centered on the axis Al, this central bore allowing the axial passage of one or more drive shafts 46 through the disk.

[0045] The bore 44 is defined at the level of an inner periphery of parts 22b, 22c. This is a region of the disc that is axially thickened, then tapers radially outwards. Each of the two parts 22b, 22c therefore has a radially internal end of maximum axial thickness, which then tapers radially outwards. A radially external fixation portion 48b, 48c, of reduced axial thickness, is also provided, projecting radially outwards. From the radially internal base of these two portions 48b, 48c, to their radially external ends, two opposing axial recesses 50b, 50c are formed, so as to jointly create a housing space for the radially internal fixation portion 34 of the first part 22a.

[0046] The two fixing portions 48b, 48c respectively have two surfaces which face each other axially, and which therefore delimit the aforementioned housing space.These two surfaces are each equipped with radial retention elements 52b, 52c for the first part 22a of the disc. The additional retention elements 52b, 52c are annular grooves, intended to receive the ridge-shaped elements 40 of the first part 22a. Of course, a reversed arrangement of the elements 40, 52b, 52c could be provided, without departing from the scope of the invention.

[0047] In addition, each radially external portion of the fixing 48b, 48c is traversed by through holes 54b, 54c, axially oriented and circumferentially spaced from each other.

[0048] In the assembled configuration of the disc as shown in Figures 2 and 5, the radially internal mounting portion 34 is axially clamped between the two radially external mounting portions 48b, 48c. Furthermore, the cumulative axial thickness of these three clamped portions 34, 48b, 48c is equal to or substantially equal to the axial thickness of the tooth support 32 of the first part 22a. Moreover, as can be seen in the figures, the portions 48b, 48c have a radially external end that extends radially to the tooth support 32 of the first part 22a. This contributes to obtaining a disk 20 of compact shape, and of relatively homogeneous axial thickness.

[0049] When the radially internal fixing portion 34 is axially clamped between the two radially external fixing portions 48b, 48c, the retention members 40, 52b, 52c cooperate in pairs so as to form radial retention means of the first part 22a, relative to the second and third parts 22b, 22c.

[0050] To achieve the aforementioned clamping, the disc 20 includes fastening means in the form of axial bolts 56, or similar fastening elements, such as rivets. Each bolt 56 passes through three aligned through holes 54b, 42, 54c, so as to apply an axial clamping force to three axially stacked fastening portions 48b, 34, 48c. Furthermore, the aforementioned radial retaining means 40, 52b, 52c can be arranged radially on either side of the bolts. This results in one or more assemblies 40, 52a, 52b radially outward relative to these bolts 56, as well as one or more assemblies 40, 52a, 52b radially inward relative to these same bolts 56.

[0051] The rotating wheel 12B is complemented by a flange 58 shown in [Fig. 2], preferably arranged upstream of the disc 20. This flange 58 allows, if necessary, a cooling airflow to be channeled along the disc. At its radially external end, it secures an axial retention ring 60 for the rotating blades 30. This ring 60, in a known manner, is positioned opposite the teeth and the recesses of the turbine disc 20, in order to prevent the blade roots from escaping axially from their recesses. In addition, the ring 60, which is pressed against the disc 20, is also housed in a hook 63 of the blade platform. This allows axial retention of the blades in both directions.

[0052] Another feature of the invention is that the flange 58 is mounted on the disc using the same bolts 56 that secure the disc parts together. Consequently, these bolts 56 also pass through holes made in the flange 58, for example at a radially internal end thereof.

[0053] Optionally, these bolts 56 can also provide the mounting of a balancing system 62 for the movable wheel 12B.

[0054] As previously stated, the two materials used are different. Preferably, the first material has superior resistance to fining compared to the second material, because the first part 22a of the disk is located closer to the primary stream than the second and third parts 22b, 22c, and is therefore more exposed to the high temperatures of the primary stream.

[0055] Furthermore, the second material exhibits superior fatigue resistance compared to the first material, because the first part 22a is less susceptible to fatigue than the first part 22a. The second and third parts 22b, 22c may be, in particular because of the presence of the central bore 44 on the latter.

[0056] By way of non-limiting examples, the two materials are preferably metallic, for example heat-resistant alloys based on Nickel or Cobalt.

[0057] Of course, various modifications can be made by a person skilled in the art to the invention which has just been described only by way of non-limiting examples, and within the limits of the scope of the annexed claims.

Claims

Demands

1. Disc (20) for a bladed rotating wheel (12B) of an aircraft turbomachine module, the disc extending circumferentially around a central longitudinal axis (Al), and being characterized in that it comprises: - a first part (22a) made of a first material, and having teeth (26) between which are defined recesses (28) for housing blades (30) of the rotating wheel; - a second part (22b) made of a second material different from the first material, the second part (22b) being assembled to the first part (22a), and defining at least partially a central disc bore (44) centered on the central longitudinal axis (Al).

2. Disc according to claim 1, characterized in that the first and second parts (22a, 22b) are annular, centered on the longitudinal central axis (Al).

3. Disc according to claim 1 or 2, characterized in that it also comprises a third part (22c), preferably made of the second material, and also defining at least partially the disc bore (44).

4. Disc according to claim 3, characterized in that the second and third parts (22b, 22c) axially enclose the first part (22a), and in that the first part (22a) comprises a radially internal fixing portion (34), axially enclosed between two radially external fixing portions (48b, 48c) belonging respectively to the second and third parts (22b, 22c), the radially internal and external fixing portions (48b, 48c) being equipped with radial retention means (40, 52b, 52c) of the first part (22a) relative to the second and third parts (22b, 22c).

5. Disc according to claim 3 or 4, characterized in that the second and third parts (22b, 22c) each have the shape of a disc with a hub and a rim, and they preferably constitute two symmetrical or substantially symmetrical parts with respect to a radial mounting interface plane between these two parts (22b, 22c).

6. Disc according to any one of claims 3 to 5, characterized in that it comprises means for attaching (56) the second part (22b) on the third part (22c), said fastening means (56) also preferably passing through the first part (22a).

7. Disc according to claim 4, characterized in that the radial retaining means (40, 52b, 52c) of the first part (22a), relative to the second and third parts (22b, 22c), comprise on the one hand, annular ridges or ribs, and, on the other hand, annular grooves configured to cooperate in a complementary manner with the annular ridges or ribs.

8. Disc according to claim 7 combined with claim 6, characterized in that the radial retaining means (40, 52b, 52c) are arranged on the one hand radially inside and, on the other hand, radially outside with respect to said fixing means (56).

9. Bladed wheel (12B) of module (8, 9) of aircraft turbomachine, comprising a disc (20) according to any one of the preceding claims, as well as movable blades (30) housed in the recesses (28) of the first part (22a), the movable wheel also comprising, preferably, a flange (58) for retaining an axial retaining ring (60) of the movable blades (30).

10. Movable wheel according to claim 9, combined with claim 6, characterized in that the means for fixing (56) the second part (22b) to the third part (22c) also pass through the flange (58) for retaining the axial retaining ring (60).

11. Aircraft turbomachine module (8, 9) comprising at least one wheel (12B) according to claim 9 or 10, the module preferably being a low pressure turbine.

12. Aircraft turbomachine (1), comprising at least one module (8, 9) according to the preceding claim.

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