AIRCRAFT TURBOMACHINE INCLUDING AN IMPROVED BEARING SUPPORT DESIGN

Abstract

The invention relates to an assembly (20) for an aircraft turbomachine, comprising a stator part (26), a first bearing (28a), a second bearing (28b) and a support piece (32) in an oil enclosure (22) delimited by an outer portion of enclosure delimitation (24) integrated into the stator part (26), the support piece (32) comprising: - a first axial end portion (34a) forming an outer ring (36) of the first bearing (28a), or supporting such a ring (36); - a second axial end portion (34b) forming an outer ring of the second bearing (28b) or supporting such a ring, a compression damping oil film (50) being arranged between the second portion (34b) and the stator part (26); - an intermediate ring (46), arranged axially between the first and second portions (34a, 34b), and forming a flexible connection as well as an oil splash protection element. Figure 2.

Description

Title of the invention: AIRCRAFT TURBOMACHINE COMPRISING AN IMPROVED BEARING SUPPORT technical field

[0001] The present invention relates to the field of aircraft turbomachinery. It relates more particularly to support and guidance systems for turbomachine engine shafts.

[0002] The invention applies to all types of aircraft, such as helicopters in particular. STATE OF PRIOR ART

[0003] Aircraft turbomachinery engine shafts are guided in rotation by bearings. Oil film compression dampers, also known as "squeeze film dampers (SFDs)," can be used with these bearings. This type of damper provides damping to flexible shafts operating at supercritical speeds. The thin film, located between the outer ring of the bearing and a portion of the stator, is not sheared, unlike in a conventional hydrodynamic bearing. This absence of shear limits oil heating and eliminates instabilities.

[0004] The role of such a damper is to dampen the vibrations of the drive shaft caused by its rotation and by the imbalance of its mechanical load, for example, unbalance. This principle of damping by compression of an oil film is, for example, known from document EP 1 650 449 AL

[0005] Typically, the outer ring of the damped bearing is integral with a flexible cage, known as a squirrel cage. This cage has openings that provide the required flexibility. It allows for greater orbital movement of the outer ring within the stator portion that houses it, thus ensuring improved efficiency of the oil film compression damper.

[0006] The flexible cage is generally surrounded by the stator part, which is likely to have a high temperature in the hot areas of the turbomachine, for example in the rear part thereof, in an oil enclosure located at the level of the turbine(s).

[0007] During operation, the rotation of components located within the oil chamber, such as the rotating elements of the bearings and the drive shafts, generates oil spray in the form of a mist. To prevent this oil from being sprayed onto the hot wall of the stator section surrounding the flexible cage, by passing through the large openings in this cage, a grid-shaped component is to be integrated. anti-projection ring inside the cage.

[0008] The incorporation of this additional component complicates the overall design of the assembly, in an already very dense environment. Furthermore, in order for the flexible cage to retain its deformation capacity, a radial clearance must be provided between the flexible cage bars and the annular grid. This clearance implies the potential passage of oil mist at the axial ends of the free annular space between the cage and the grid, with a risk of oil being projected radially outwards against the hot wall of the stator section, passing through the large openings in the flexible cage.

[0009] Due to oil splashes observed on the hot part of the stator, and despite the presence of the annular grid intended to limit these splashes, the impacts cause forced convection-type heat exchange. This results in significant local increases in oil temperature. These temperature increases negatively impact the oil's heat balance, necessitating oversizing of the heat exchangers designed to cool it. Furthermore, causing excessively high local oil temperatures generates rapid thermal oxidation of the oil, leading to problems in addressing the risks of coking. Description of the invention

[0010] To address at least partially the problems mentioned above relating to prior art, the invention first relates to an assembly for an aircraft turbomachine, comprising a stator portion, a second bearing, and a support piece, the second bearing and the support piece being arranged in an oil chamber radially delimited outwards by an external portion of the chamber boundary integrated into the stator portion, the support piece comprising:

[0011] - a first axial end portion preferably intended to form a ring external of a first bearing, or to support such a ring;

[0012] - a portion for attaching the support piece to the stator part, the portion of fixing being connected to the first portion of axial end;

[0013] - a second axial end portion opposite the first end portion axial, and forming an outer ring of the second bearing or supporting such a ring, a compression-damping oil film being arranged between an outer surface of the second axial end portion and a corresponding surface of the stator portion; and

[0014] - an intermediate ring, arranged axially between the first and second portions axial end, and forming a flexible connection between these two same axial end portions.

[0015] According to the invention, the intermediate ring also forms a protective element oil splash guard, configured to prevent oil from splashing from inside this intermediate ring against the outer portion of the enclosure boundary.

[0016] Thus, the intermediate ring performs two functions through this same component. The first function is the flexible mechanical support of the second axial end portion associated with the damping oil film, while the second thermal function aims to reduce the oil temperature, preventing it from impacting the outer portion of the enclosure boundary, which constitutes a potentially hot zone of the turbomachine. Unlike prior art embodiments in which two separate parts were required to perform these two functions, the second of which was not even optimally implemented, the invention advantageously provides for fusing these two parts into one, with a much finer perforation of the ring to prevent unwanted projections of oil particles.

[0017] The overall design is simplified, and the thermal function is improved.

[0018] The invention also has at least one of the following optional features, taken individually or in combination.

[0019] Preferably, the oil splash protection element is traversed by orifices, each made so that its smallest dimension is less than 5 mm.

[0020] According to a first preferred embodiment of the invention, the oil splash guard has orifices arranged in succession along a circumferential direction of the ring, as well as along an axial direction of the ring, these orifices preferably being arranged in annular rows of axially successive orifices. In this configuration, the oil splash guard thus takes on a general grid shape. The orifices are preferably circular, oval, or oblong, or any other shape deemed appropriate.

[0021] According to a second preferred embodiment of the invention, the oil splash protection element has orifices following one another only in a circumferential direction of the ring, each orifice having a slot shape extending from one axial end to the other of the intermediate ring.

[0022] Whatever the embodiment envisaged, the protective element has openings as well as solid portions delimiting these openings, and the ratio between the cumulative area of ​​the openings, and the total area of ​​the protective element, is between 0.2 and 0.5. It can for example be in the order of 0.3 to 0.4.

[0023] Preferably, the portion for fixing the support piece on the stator part takes the form of a fixing flange, or a plurality of fixing lugs distributed circumferentially around the first axial end portion of the support piece.

[0024] Preferably, the outer portion of the enclosure delimitation, integrated into the stator part, has an oil recovery orifice in its lower part.

[0025] Preferably, the support piece specific to the invention and described above is preferably made in one piece, for example by any additive manufacturing technique offering the possibility of a complex geometry. However, other manufacturing techniques are conceivable, for example by making the support piece in several portions fixedly attached to one another, and / or by providing machining / drilling to create the openings for the protective element.

[0026] The invention also relates to an aircraft turbomachine comprising at least one such assembly, preferably arranged downstream of a combustion chamber of the turbomachine.

[0027] Preferably, the turbomachine comprises a first drive shaft guided by the first bearing, as well as a second drive shaft guided by the second bearing, the first and second shafts being concentric.

[0028] Other advantages and features of the invention will appear in the detailed, non-limiting description below. Brief description of the drawings

[0029] This description will be made with reference to the attached drawings, among which:

[0030] [Fig-1] represents a schematic axial cross-sectional view of a turbomachine aircraft according to the invention;

[0031] [Fig.2] represents a schematic enlarged axial cross-sectional view of an assembly equipping the rear part of the turbomachine shown in the previous figure, the assembly being in the form of a first preferred embodiment of the invention;

[0032] [Fig.3] is an enlarged perspective view of the support piece equipping the assembly shown in the previous figure;

[0033] [Fig.4] is an axial cross-sectional view of the support piece shown in the previous figure;

[0034] [Fig. 5] is a perspective view similar to that of [Fig. 3], with the part presented in the form of a second preferred embodiment of the invention; and

[0035] [Fig.6] is an axial cross-sectional view of the support piece shown in the previous figure. Description of the implementation methods

[0036] With reference first to [Fig. 1], a preferred embodiment of the invention is schematically represented, taking the form of an aircraft turbomachine 1, here a helicopter.

[0037] The turbomachine 1 includes a main transmission gearbox 2, intended to drive a rotating receiver (not shown), namely helicopter blades in this case. The main transmission housing is connected to a gas generator 4 of the turbomachine, comprising, from upstream to downstream along a main flow direction 6 of the gases through the turbomachine, a centrifugal compressor 8, possibly two-stage, a combustion chamber 10, a high-pressure turbine 12, and a free turbine 14, also called a low-pressure turbine.

[0038] The compressor 8 and the turbine 12 are connected by a drive shaft 16a, corresponding to a high-pressure shaft, while the free turbine 14 is connected to the downstream end of a free turbine shaft 16b, corresponding to a low-pressure shaft. The upstream end of this shaft 16b is connected to the main transmission 2, into which it drives the transmission. Hereafter, the shafts 16a and 16b will be referred to as the first and second drive shafts. They are concentric, centered on a longitudinal central axis 18 of the gas generator 4. Preferably, the first shaft 16a is located around the second shaft 16b.

[0039] The invention relates more specifically to an assembly 20 located in the rear part of the turbomachine, arranged wholly or partly downstream of the combustion chamber 10. The assembly 20 is preferably located between the two turbines 12, 14, and defines an oil chamber 22 which is radially delimited outwards by an outer portion of the chamber boundary 24, this portion 24 belonging to a stator portion 26 of the assembly 20. The stator portion 26 is arranged axially between the two turbines, such that its outer portion of the chamber boundary 24 corresponds to an inter-turbine housing. The stator portion 26, connected to the turbine housing 25, takes the form of a casing with its outer annular portion 24 surrounding the shafts 16a, 16b.

[0040] Within the oil chamber 22, the assembly comprises a first bearing 28a and a second bearing 28b located downstream of the latter. The first bearing 28a, optional in the implementation of the general principle of the invention, but retained in this preferred embodiment, is a roller bearing guiding the rotation of the first drive shaft 16a, while the second bearing 28b is a ball bearing guiding the rotation of the second drive shaft 16b. In the invention, it is this second bearing 28b that is associated with a squeeze film damper, as will be detailed below.

[0041] As can be most clearly seen in [Fig. 2], upstream of the first bearing 28a, the oil chamber 22 is axially closed by a first seal 30a, for example of the labyrinth type, interposed between an upstream end of the outer annular portion 24 of the stator part 26, and an outer surface of the first shaft 16a. Similarly, downstream of the second bearing 28b, the oil chamber 22 is axially closed by a second seal 30b, for example of the labyrinth type, interposed between a downstream end of the outer annular portion 24 of the stator part 26, and an outer surface of the second shaft 16b.

[0042] Each of the two drive shafts 16a, 16b therefore penetrates axially into the stator part 24 and into the oil chamber 22, the latter also containing a support piece 32 specific to the invention, and of which a first embodiment is shown in figures 1 to 4.

[0043] This support piece 32 is preferably made in one piece rather than by assembling several components, although the latter remains a possibility. The one-piece part is produced, for example, by any additive manufacturing technique, preferably in a metallic material, although other conventional manufacturing techniques remain possible.

[0044] The support piece 32 is generally annular and cylindrical, centered on the axis 18. It comprises the various elements described below:

[0045] - a first axial end portion 34a forming a support housing a ring external 36 of the first bearing 28a, the inner ring of this bearing being carried externally by the first shaft 16a;

[0046] - a fixing portion 40 of the support piece 32, on the annular portion ex external 24 of the stator part 26. The fixing portion 40 is connected to the first axial end portion 34a, forming a plurality of fixing lugs 40 distributed circumferentially around a downstream end of the first axial end portion 34a, and cooperating preferentially with stator lugs (not shown) projecting radially inside the enclosure 22 from the external enclosure delimitation portion 24;

[0047] - a second axial end portion 34b opposite the first portion axial end 34a, and forming an outer ring of the second bearing 28b, with an inner race 42 for the bearing balls. The inner ring 44 of the second bearing is carried externally by the second drive shaft 16b; and

[0048] - an intermediate ring 46 specific to the invention, arranged axially between the first and second axial end portions 34a, 34b, and forming a flexible connection between them.

[0049] The ring 46, which forms a flexible cage, is indeed finely perforated so as to provide the flexibility required for the mechanical retention of the second axial end portion 34b. This second portion 34b, corresponding to the outer ring of the second bearing 28b, is in fact damped by a compression damping oil film 50, arranged in an annular space between an outer surface of this second portion 34b, and a corresponding surface of the outer annular portion 24 of the stator part 26.

[0050] The mechanical flexibility of the ring 46, combined with the damping oil film 50 effectively allows for a greater orbit of the outer ring 34b within the stator portion 26 that houses it, thus ensuring improved efficiency of the oil film compression damper. The latter is confined between two axially spaced annular sealing tabs 51, which are radially inwardly directed by the outer annular portion 24 of the stator portion 26.

[0051] In addition, it is noted that the outer annular portion 24 of the stator part 26 is also traversed by one or more oil injectors 53 in the enclosure, these injectors approaching, for example, as closely as possible to the shafts 16a, 16b.

[0052] In addition to providing the required flexibility to the ring 46 intended to center the outer ring 34b of the second bearing 28b, this ring has the particularity of incorporating a second function of the thermal protection type for the oil. Indeed, the intermediate ring 46 also forms an oil splash protection element, configured to prevent oil from being projected from the inside of this intermediate ring against the inner surface of the outer portion of the enclosure boundary 24. In other words, the ring 46 has a certain porosity and is sufficiently finely perforated so that, during operation, the particles of oil mist projected radially outwards by the rotation of the shafts 16a, 16b and / or that of the rolling elements of the bearings 28a, 28b, are slowed down / stopped by this protective element 46.The oil projected against the latter can then obviously pass through this element, but its practical purpose is to reduce the oil's speed or even stop it completely to prevent it from impacting the outer portion of the enclosure boundary 24. After passing through the protective element, the oil can nevertheless come into contact with part of the inner surface of the outer portion of the enclosure boundary 24, for example, after running down the protective element and then falling by gravity onto or near a lower part of the stator housing containing one or more oil recovery ports 54. The contact of the oil with the inner surface of the outer portion of the enclosure boundary 24 therefore does not occur through projection of this oil from inside the ring 46, but is above all partial because it only concerns a part of the aforementioned inner surface.Due to this partial contact, without oil splashing onto the crankcase, the increase in oil temperature by convection is advantageously contained. The risks of coking are thus reduced, and the oil cooling requirements are more moderate, necessitating less bulky heat exchangers.

[0053] In the first preferred embodiment of the invention, the intermediate ring 46 is finely perforated by having orifices 56 arranged in succession along a circumferential direction "C" of the ring, as well as along an axial direction "A" of the ring. Thus, the orifices 56 are preferably arranged in annular rows of orifices. following one another axially, thus forming a grid-like, matrix-shaped splash guard. A staggered arrangement is also possible. Here, the orifices 56 of the splash guard 46 are circular, but other shapes can be adopted without departing from the scope of the invention. To ensure the desired splash guard function while offering the required flexibility, the orifices 56 are small and densely packed. Each orifice 56 is designed so that its smallest dimension "Dmin," that is, its diameter in the case of a circular shape where its dimension remains the same in all directions, is less than 5 mm. For example, this smallest dimension Dmin can be on the order of 2 mm. Furthermore, the density of orifices 56 within the element 46 can be on the order of four to five orifices / cm².

[0054] A second embodiment is shown in Figures 5 and 6. It differs from the previous embodiment in that the orifices 56 of the protective element are arranged here only in the circumferential direction "C" of the ring, and no longer also in the axial direction "A". Thus, each orifice 56 has a slot shape extending continuously from one axial end to the other of the ring 46 in the direction "A3", preferably having a straight and purely axial shape as in Figures 5 and 6, or having, for example, a circumferential component.

[0055] These orifices 56 are each made so that its smallest dimension “Dmin”, i.e. here the slit width along the circumferential direction “C”, is less than 5 mm. It can for example be on the order of 1.5 mm.

[0056] To provide the required flexibility while maintaining the small orifice width indicated above and necessary for the oil-splash protection function, the solid, bar-shaped portions 58 remaining within the element 46 are preferably of small cross-section. Furthermore, the ratio between the cumulative surface area of ​​the orifices 56 of this element and the total surface area of ​​the latter, including the orifices 56 and the bars 58, is preferably between 0.2 and 0.5, for example, being set at 0.4 in this preferred embodiment. This range of values ​​concerning the proportion of the cross-sectional area of ​​the protective element is also applicable to the first preferred embodiment described above, in which the ratio is, for example, set at 0.3. By way of example, the cross-sectional area provided by all the orifices 56 is on the order of 2,000 mm².

[0057] 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 whose scope is delimited by the appended claims. In particular, the turbomachine described above is intended for a helicopter, but it could alternatively be designed for the propulsion of an airplane.

Claims

Claims

1. Assembly (20) for an aircraft turbomachine, comprising a stator part (26), a second bearing (28b), and a support part (32), the second bearing (28b) and the support part being arranged in an oil chamber (22) delimited radially outwardly by an outer chamber delimiting portion (24) integrated into the stator part (26), the support part (32) comprising: - a first axial end portion (34a), preferably intended to form an outer ring (36) of a first bearing (28a), or to support such a ring (36); - a fixing portion (40) of the support part on the stator part (26), the fixing portion (40) being connected to the first axial end portion (34a);- a second axial end portion (34b) opposite the first axial end portion (34a), and forming an outer ring of the second bearing (28b) or supporting such a ring, a compression damping oil film (50) being arranged between an outer surface of the second axial end portion (34b) and a corresponding surface of the stator part (26); and - an intermediate ring (46), arranged axially between the first and second axial end portions (34a, 34b), and forming a flexible connection between these same two axial end portions, characterized in that the intermediate ring (46) further forms an oil splash protection element, configured to prevent oil from being splashed from the inside of this intermediate ring (46), against the outer enclosure delimiting portion (24).;

2. Assembly according to claim 1, characterized in that the oil splash protection element (46) is crossed by orifices (56), each made so that its smallest dimension (Dmin) is less than 5 mm.

3. Assembly according to claim 1 or 2, characterized in that the oil splash protection element (46) has orifices (56) succeeding one another in a circumferential direction (C) of the ring, as well as in an axial direction (A) thereof, these orifices (56) being preferentially arranged in annular rows of orifices succeeding one another axially.

4. Assembly according to claim 3, characterized in that the orifices (56) of the protective element are circular, oval or oblong in shape.

5. Assembly according to claim 1 or 2, characterized in that the oil splash protection element (46) has orifices (56) succeeding one another only in a circumferential direction (C) of the ring, each orifice (56) having a slot shape extending from one axial end to the other of the intermediate ring.

6. Assembly according to any one of the preceding claims, characterized in that the oil splash protection element (46) has orifices (56) as well as solid portions (58) delimiting these orifices, and in that the ratio between the cumulative surface area of the orifices and the total surface area of the protection element is between 0.2 and 0.

5.

7. Assembly according to any one of the preceding claims, characterized in that the fixing portion (40) of the support piece (32) on the stator part (26) takes the form of a fixing flange, or of a plurality of fixing ears distributed circumferentially around the first axial end portion (34a) of the support piece (32).

8. Assembly according to any one of the preceding claims, characterized in that the outer enclosure delimiting portion (24), integrated into the stator part (26), has an oil recovery orifice (54) in the lower part.

9. Aircraft turbomachine (1) comprising at least one assembly (20) according to any one of the preceding claims.

10. Turbomachine according to claim 9, characterized in that it comprises a first drive shaft (16a) guided by the first bearing (28a), as well as a second drive shaft (16b) guided by the second bearing (28b), the first and second shafts (28a, 28b) being concentric.

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