Damping device for rotor wheel of an aircraft turbomachine, rotor wheel of an aircraft turbomachine, aircraft turbomachine and method of manufacturing a damping device

Abstract

The invention relates to a damping device (40) for a rotor wheel (8) of a turbine engine (1) of an aircraft, which damping device comprises two support elements (41, 42) and a resilient member (43) arranged between the two support elements (41, 42) so as to allow a relative movement of the support elements (41, 42) towards one another under the action of external forces exerted on these support elements (41, 42) and compression of the resilient member (43). The device (40) is intended to be arranged between two blades (21) of said rotor wheel (8) that follow one another circumferentially such that one of the support elements (41) is opposite a shank (25) of one of these blades (21) and the other support element (42) is opposite a shank (25) of the other of said blades (21). Each of the support elements (41, 42) forms a casing that defines a hollow space in which a relevant part of the resilient member (43) is housed. The invention also relates to said rotor wheel (8) for an aircraft turbine engine (1), said turbine engine (1) of an aircraft, and a method for manufacturing said damping device (40).

Description

Technical field

[0001] The invention relates to the field of turbomachines for aircraft propulsion systems. The invention relates to a damping device for a moving wheel of an aircraft turbomachine, a moving wheel of an aircraft turbomachine, an aircraft turbomachine. It also relates to a method for manufacturing a damping device for a moving wheel of an aircraft turbomachine. State of the prior art

[0002] The rotor of a conventional aeronautical turbomachine comprises moving wheels each formed by a disc and a plurality of blades which follow one another circumferentially. Each blade comprises a root received in a respective cell formed at the periphery of the disc.

[0003] In order to dampen the vibration responses of such a wheel during operation of the turbomachine, it is known to arrange sheet metal members between the disc and the blades so as to dissipate the vibration energy by dry friction between these members and the moving wheel.

[0004] Such a damping member is generally housed in a cavity delimited circumferentially by the stilts of two adjacent blades, radially outwards by the platforms of these blades and radially inwards by a tooth of the disc.

[0005] Under the action of the rotation of the moving wheel around the radial axis, such a damping member can be caused to pivot around the radial axis in such a way as to be pressed axially upstream on one of the blade stilts delimiting the cavity and axially downstream on the other blade stilt delimiting the cavity, which tends to damage the stilts.

[0006] FR 2 663 686 A1 and FR 1 579 923 A each disclose a damping device for a moving wheel of an aircraft turbomachine according to the state of the prior art. Statement of the invention

[0007] The invention aims to provide a damping device making it possible to reduce the wear of surrounding parts.

[0008] To this end, the invention relates to a damping device for a moving wheel of an aircraft turbomachine, according to the characteristics of claim 1.

[0009] The elastic member allows the shock absorber to deform by allowing relative movements of the support elements and dissipating part of the forces to which it is subjected.

[0010] The damping device of the invention thus makes it possible to reduce the tangential forces applied to the stilts, or more generally to the parts with which the support elements come into contact, and thus to reduce their wear, while ensuring good damping performance.

[0011] Preferably, each envelope comprises a free edge, an inner surface delimiting the hollow space of this envelope and an outer surface, the inner surface and the outer surface of each envelope being delimited by the free edge of this envelope, the device being configured so that the free edges of the two envelopes are opposite each other.

[0012] The free edges of the two envelopes define between them an interval making it possible to prevent them from coming into contact with each other, at least when the device is not subjected to external stress and preferably in all phases of operation of the turbomachine.

[0013] Such an arrangement of the free edges makes it possible, on the one hand, to deform the device by compression of the elastic member under the action of a force exerted on the envelopes and, on the other hand, to prevent the free edges of the envelopes from coming into contact with the stilts delimiting the cavity receiving the device.

[0014] The outer surface of each of the envelopes is preferably smooth.

[0015] In other words, the outer surface of each of the envelopes is preferably free of edges or protruding parts.

[0016] In one embodiment, the outer surface of each of the envelopes comprises a lateral portion intended to be arranged opposite said stilt of a respective one of said blades.

[0017] In one embodiment, the outer surface of each of the envelopes comprises a lower part intended to be arranged opposite a tooth of a disc of said mobile wheel.

[0018] In an embodiment in which the outer surface of each of the envelopes comprises a side portion and a bottom portion, the bottom portion is preferably connected to the side portion by a rounding.

[0019] In one embodiment, the outer surface of each of the envelopes comprises an upper portion intended to be arranged opposite a platform of a respective one of said blades.

[0020] In an embodiment in which the outer surface of each of the envelopes comprises a side portion and a top portion, the top portion is preferably connected to the side portion by a rounding.

[0021] In one embodiment, the elastic member comprises a first material and the support elements comprise a second material different from the first material.

[0022] The first material may be an elastomeric material preferably comprising polyurethane or fluoroethane.

[0023] The second material may be a metal alloy preferably comprising nickel.

[0024] Such a metal alloy allows to obtain a high coefficient of friction with the moving wheel, which increases the efficiency of vibration energy dissipation during friction.

[0025] The invention also relates to a mobile wheel for an aircraft turbomachine, extending around a longitudinal axis and comprising a disc and blades, the disc forming teeth circumferentially defining between them cells, the blades each comprising a platform, a stilt and a root received in a respective one of the cells, the wheel forming a cavity circumferentially delimited by the stilts of two of said blades which follow one another circumferentially, the cavity being delimited radially inwards by one of the teeth extending circumferentially between the roots of these two blades and radially outwards by the platforms of these two blades, the wheel comprising a damping device as described above housed in the cavity.

[0026] The invention also relates to a turbomachine for aircraft, comprising a mobile wheel as described above.

[0027] According to another aspect, the invention relates to a method of manufacturing a damping device as described above.

[0028] In one embodiment, this method comprises a step of stamping two sheets so as to form said envelopes.

[0029] In one embodiment, the method comprises a step of assembling the elastic member and the two envelopes or support elements.

[0030] Preferably, the assembly step comprises a cold bonding or hot bonding operation.

[0031] Such assembly methods are simple and inexpensive.

[0032] Other advantages and characteristics of the invention will appear on reading the detailed, non-limiting description which follows. Brief description of the drawings

[0033] The following detailed description refers to the attached drawings in which: [ Fig. 1] is a schematic axial sectional view of a twin-spool, twin-flow turbojet; [ Fig. 2 ] is a schematic perspective view of an angular sector of a moving wheel according to the invention; [ Fig. 3 ] is a schematic perspective and cross-sectional view of part of the wheel of the Figure 2 ; [ Fig. 4 ] is a schematic view of part of the wheel of the Figure 2 , including a shock absorber according to the invention; [ Fig. 5 ] is a schematic perspective and exploded view of the shock absorber of the Figure 4 ; [ Fig. 6 ] is a schematic perspective view of the shock absorber of the Figure 5 ; [ Fig. 7 ] is a schematic perspective and cross-sectional view of the shock absorber of the Figure 6 ; [ Fig. 8 ] is a schematic perspective and cross-sectional view of part of the wheel of the Figure 2 , including the shock absorber of the Figure 5 ; [ Fig. 9] is a schematic perspective view of part of the wheel of the Figure 2 , including the shock absorber of the Figure 5 . Detailed description of embodiments

[0034] It is represented on the Figure 1 a twin-spool, low bypass turbojet 1, intended for aircraft propulsion.

[0035] The turbojet 1 has a longitudinal central axis A1 around which its various components extend.

[0036] In the present description, the terms “upstream” and “downstream” are defined with reference to a direction S1 of main flow of the gases in the turbojet 1 along the axis A1.

[0037] THE figures 1 to 9 include a reference defining an axial or longitudinal direction D1 corresponding to the direction of the axis A1, a radial direction D2 and a tangential or circumferential direction D3.

[0038] The turbojet 1 comprises, from upstream to downstream, a low pressure compressor 2, a high pressure compressor 3, a combustion chamber 4, a high pressure turbine 5 and a low pressure turbine 6 constituting a gas generator.

[0039] Low pressure compressor 2 and low pressure turbine 6 form a low pressure body. High pressure compressor 3 and high pressure turbine 5 form a high pressure body.

[0040] In a manner known per se, the compressors 2 and 3 and the turbines 5 and 6 each comprise a rotor and a stator forming one or more stages. Generally, a compressor stage comprises, from upstream to downstream, a bladed moving wheel which draws in and accelerates an air flow and a rectifier provided to straighten the flow thus accelerated by increasing its pressure. A turbine stage, for its part, comprises, from upstream to downstream, a distributor and a bladed moving wheel, the distributor being configured to accelerate the air flow towards the wheel in order to drive it in rotation.

[0041] In the example of the Figure 1 , the low pressure compressor 2 comprises seven stages, the high pressure compressor 3 comprises four, the high pressure turbine 5 comprises one and the low pressure turbine 6 comprises two. Of course, each of the modules of the gas generator can have a different number of stages without departing from the scope of the invention.

[0042] On the Figure 1 , the moving wheel and the rectifier of the first stage of the low pressure compressor 2 are referred to by the references 8 and 9 respectively. The moving wheel of the second stage of this compressor 2 is referred to by the reference 10 and that of the third stage is referred to by the reference 11.

[0043] The following description relates by way of non-limiting example to the moving wheel 8 of the first stage of the low pressure compressor 2 of the Figure 1 .

[0044] Of course, the wheel described below can form another wheel of the compressor 2, for example the wheel 10 or 11, or of another module of the turbojet 1 of the Figure 1 , for example from the high pressure compressor 3, or even a wheel of a turbomachine different from that of the Figure 1 The following description applies by analogy to these different applications.

[0045] THE figures 2 And 3show an angular sector of the moving wheel 8.

[0046] As is known per se, the wheel 8 comprises a disc 20 and blades 21 - two of which are visible on the figures 2 And 3 - carried by the disc 20 while being arranged next to each other in the circumferential direction D3.

[0047] For this purpose, the disc 20 has on its periphery teeth 22 which circumferentially define between them blade housing cells. In this example, each cell extends in a direction substantially parallel to the axis A1, so as to form an opening passing through the disc 20 from upstream to downstream.

[0048] Each blade 21 comprises, radially from the inside to the outside, a foot 24, a stilt 25, a platform 26 and a blade 27 forming the aerodynamic part of the blade 21 (see Figure 3). In this example, the foot 24 of each blade 21 has an external shape called “fir tree” or “bulb”, allowing its insertion into one of the cells of the disc which has a corresponding shape.

[0049] Each blade 21 is thus mounted on the disc 20 by inserting its foot 24 into one of the respective cells.

[0050] With reference to the two blades 21 visible on the figures 2 And 3 , which follow one another circumferentially, these delimit with the disc 20 a cavity. This cavity is delimited circumferentially by the stilts 25 of these two blades 21, radially inwards by the tooth 22 extending circumferentially between the feet 24 of these two blades 21 and radially outwards by the platforms 26 of these two blades 21.

[0051] Each pair of blades 21 which follow one another circumferentially delimit a similar cavity. The following description relates to the cavity extending between the two blades 21 shown in the figures 2 And 3 and applies by analogy to each of the other cavities.

[0052] In reference to the Figures 3 and 4 , each tooth 22 of the disc 20 forms an ear 32 extending radially outwardly at a downstream end of the tooth 22 so as to form a downstream surface of the disc 20.

[0053] There Figure 4 shows a retaining ring 33 resting on the downstream surface formed by the teeth 22 of the disc 20 of the wheel 8.

[0054] The ring 33 is fixed to the disc by connecting means 34 forming in this example bolts which each pass through an orifice made in one respective of the ears 32.

[0055] The ring 33 has a radial dimension allowing it to axially close downstream the blade housing cells formed by the disc 20 of the wheel 8, thus forming a downstream axial stop for the blades 21.

[0056] In this example, the ring 33 is formed by an upstream end of a ferrule secured to the disc of the wheel 10 of the second stage of the compressor 2 of the Figure 1 .

[0057] Another annular stop member (not shown) is in this example mounted upstream of the disc 20 so as to form an upstream axial stop for the blades 21.

[0058] The invention relates more specifically to a damping device 40, also called a damper, designed to be housed in the aforementioned cavity so as to dampen the vibration responses of the wheel 8 during operation of the turbojet 1.

[0059] THE figures 5 to 7 show a shock absorber 40 in accordance with the invention, the figures 4 , 8 And9 illustrating its relative position in the cavity, relative to the other parts of the wheel 8.

[0060] The D1-D2-D3 reference framework of figures 5 to 7 indicates the relative position of the shock absorber 40 when it is thus housed in the cavity.

[0061] In reference to the Figure 5 , the shock absorber 40 comprises three parts 41, 42 and 43, the parts 41 and 42 forming support elements, the part 43 forming an elastic member.

[0062] In this example, the support elements 41 and 42 are each formed by a nickel-based metal sheet having a thickness less than or equal to 1 mm which is stamped so as to form a hollow space.

[0063] The support element 41 comprises an inner surface 50 and an outer surface 51 which define between them the thickness of the sheet metal forming this support element 41.

[0064] The support element 41 forms a free edge 52 which delimits the inner 50 and outer 51 surfaces and which defines a closed curve.

[0065] From a geometric point of view, the inner surfaces 50 and outer surfaces 51 are non-developable surfaces.

[0066] The inner surface 50 of the support element 41 delimits said hollow space formed by this part.

[0067] In reference to the figures 5 to 7 , the support element 41 forms a lateral part 54, a lower part 55, an upper part 56, a downstream part 57 and an upstream part 58. The lower 55, upper 56, downstream 57 and upstream 58 parts are each delimited by a respective part of the free edge 52.

[0068] The support element 41 is shaped so that the outer surface 51 is smooth.

[0069] In particular, the lower 55, upper 56, downstream 57 and upstream 58 parts are each connected on the one hand to each other and on the other hand to the lateral part 54 by smooth transitions forming rounded shapes on the outer surface 51.

[0070] In this example, the lower 55 and upper 56 parts extend opposite each other and substantially perpendicularly to the lateral part.

[0071] The free edge 52 of the support element 41 extends in a plane which is in this case parallel to the directions D1 and D2.

[0072] The support elements 41 and 42 are symmetrical with respect to each other so that what has just been described applies by analogy to the support element 42.

[0073] Concerning the elastic member 43, this is in this example made from a material, for example based on polyurethane or fluoroethane.

[0074] As it appears from the figures 5 to 7, the elastic member 43 is arranged between the support elements 41 and 42 so that a part of the elastic member 43 is housed in the hollow space formed by the support element 41 and another part of the elastic member 43 is housed in the hollow space formed by the support element 42, so that the free edge 52 of the support element 41 is opposite the free edge of the support element 42, defining between them an interval X1 (see Figure 7 ).

[0075] Each of the support elements 41 and 42 thus forms an envelope or half-shell, defining a hollow space in which a respective part of the elastic member 43 is housed.

[0076] In this example, the elastic member 43 is fixed to the inner surface 50 of the lateral part 54 of each of the support elements 41 and 42, using a cold bonding or hot adhesion type process.

[0077] The interval X1 between the free edges 52 of the support elements 41 and 42 allows a relative movement of these parts towards each other under the action of external forces exerted on their lateral part 54 and a corresponding compression of the elastic member 43.

[0078] In reference to the figure 8 , the shock absorber 40 is housed in the aforementioned cavity of the wheel 8 so that: the lateral part 54 of the support element 41 is arranged opposite the tang 25 of a first of the blades 21 delimiting this cavity and that the lateral part of the support element 42 extends opposite the tang 25 of a second of the blades 21 delimiting this cavity, the lower part 55 of each of the support elements 41 and 42 is arranged opposite the tooth 22 delimiting this cavity (see also figures 4 And 9), the upper part 56 of the support element 41 is arranged opposite the platform 26 of said first blade 21 and the upper part of the support element 42 is arranged opposite the platform 26 of said second blade 21.

[0079] The damper 40 of the invention thus makes it possible to reduce the wear of the stilts 25 with which it comes into contact during operation of the turbojet 1.

[0080] Of course, such a damper 40 can be arranged in each of the cavities formed by the different pairs of adjacent blades 21 of the wheel 8 or of another moving wheel.

[0081] In addition, the support elements 41 and 42 and the elastic member 43 may have a geometry different from that described above, in particular depending on the geometry of the cavity in which the shock absorber 40 is intended to be housed.

Claims

1. Damping device (40) for the impeller (8) of an aircraft turbomachine (1), comprising two support elements (41, 42) and an elastic member (43) arranged between the two support elements (41, 42) so as to allow a relative movement of the support elements (41, 42) towards one another under the action of external forces exerted on these support elements (41, 42) and of compression of the elastic member (43), this device (40) being intended to be arranged between two blades (21) of said impeller (8) which are arranged in circumferential succession such that one of the support elements (41) is opposite a shank (25) of one of these blades (21) and the other support element (42) is opposite a shank (25) of the other of said blades (21), characterized in that each of the support elements (41, 42) forms a shell defining a hollow space in which a respective part of the elastic member (43) is housed.

2. Device (40) according to claim 1, wherein each shell (41, 42) comprises a free edge (52), an inner surface (50) delimiting the hollow space of this shell (41, 42) and an outer surface (51), the inner surface (50) and the outer surface (51) of each shell (41, 42) being delimited by the free edge (52) of this shell, the device (40) being designed such that the free edges (52) of the two shells (41, 42) are facing one another.

3. Device (40) according to claim 2, wherein the outer surface (51) of each of the shells (41, 42) is smooth.

4. Device (40) according to either one of claims 2 or 3, wherein the outer surface (51) of each of the shells (41, 42) comprises a side part (54) intended to be arranged opposite said shank (25) of a respective one of said blades (21), a lower part (55) intended to be arranged opposite a tooth (22) of a disc (20) of said impeller (8) and an upper part (56) intended to be arranged opposite a platform (26) of a respective one of said blades (21), the lower (55) and upper (56) parts each being connected to the side part (54) by a curve.

5. Device (40) according to any one of claims 1 to 4, wherein the elastic member (43) comprises a first material, such as an elastomeric material preferably comprising polyurethane or fluoroethane, and the support elements (41, 42) comprise a second material, such as a metal alloy preferably comprising nickel.

6. Impeller (8) for an aircraft turbomachine (1), extending about a longitudinal axis (A1) and comprising a disc (20) and blades (21), the disc (20) forming teeth (22) defining recesses between them circumferentially, the blades (21) each comprising a platform (26), a shank (25) and a root (24) received in a respective one of the recesses, the impeller (8) forming a cavity delimited circumferentially by the shanks (25) of two of said blades (21) which are arranged in circumferential succession, the cavity being delimited radially towards the inside by one of the teeth (22) extending circumferentially between the roots (24) of these two blades (21) and radially towards the outside by the platforms (26) of these two blades (21), characterized in that it comprises a damping device (40) according to any one of claims 1 to 5 housed in the cavity.

7. Aircraft turbomachine (1), comprising an impeller (8) according to claim 6.

8. Method for manufacturing a damping device (40) according to any one of claims 1 to 5, this method comprising a step of stamping two metal sheets so as to form the shells (41, 42) and a step of assembling the elastic member (43) and the two shells.

9. Method according to claim 8, wherein the assembly step comprises a cold or hot bonding operation.

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