Aircraft turbomachine comprising a bearing support having an improved design
Patent Information
- Application Number
- EP2023719824
- Authority / Receiving Office
- EP · EP
- Patent Type
- Applications
- Current Assignee / Owner
- Priority Date
- 2022-04-04
- Filing Date
- 2023-03-31
- Publication Date
- 2025-02-12
AI Technical Summary
Aircraft turbomachine engine shafts experience oil mist projections onto hot stator parts due to the design of existing bearing support systems, leading to increased oil temperatures, thermal imbalance, and potential coking issues, which complicate the design and require oversized cooling exchangers.
An integrated support assembly with a perforated intermediate ring that serves as both a flexible mechanical support and an anti-oil projection protection element, preventing oil mist from reaching the hot stator parts by slowing down or stopping oil particles through finely perforated orifices, thereby reducing temperature increases and maintaining mechanical flexibility.
The solution simplifies the design, reduces oil temperature impacts on the stator parts, minimizes thermal oxidation risks, and reduces the need for bulky cooling exchangers, enhancing the thermal balance and mechanical efficiency of the turbomachine.
Smart Images
Figure 1.1
Abstract
Description
[0001] DESCRIPTION
[0002] TITLE: AIRCRAFT TURBOMACHINE COMPRISING A BEARING SUPPORT WITH AN IMPROVED DESIGN
[0003] TECHNICAL FIELD
[0004] The present invention relates to the field of aircraft turbomachines. It relates more particularly to support and guide systems for turbomachine engine shafts.
[0005] The invention applies to any type of aircraft, such as helicopters in particular.
[0006] STATE OF PRIOR ART
[0007] Aircraft turbomachine engine shafts are guided in rotation by bearings. These bearings can be combined with oil film compression dampers, also called "squeeze film dampers (SFDs). 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 part of the stator, is not sheared unlike a conventional hydrodynamic bearing. This absence of shear has the effect of limiting oil heating and eliminating instabilities.
[0008] The role of such a damper is to dampen the vibrations of the motor shaft caused by its rotation and by the imbalance of its mechanical loading, for example the unbalance. This principle of damping by compression of an oil film is for example known from document EP 1 650 449 Al.
[0009] Typically, the outer ring of the damped bearing is secured to a flexible cage, known as a squirrel cage. This cage is crossed by openings that provide the required flexibility. It allows for greater orbiting of the outer ring in the part of the stator that houses it, and thus ensures greater efficiency of the oil film compression damper.
[0010] 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 of the latter, in an oil enclosure located at the level of the turbine(s).
[0011] During operation, the rotation of the components located in the oil enclosure, such as the rotating elements of the bearings and the motor shafts, generates oil projections in the form of mist. In order to prevent this oil from being projected onto the hot wall of the stator part surrounding the flexible cage, passing through the large openings of this cage, it is planned to integrate a part in the form of an annular anti-projection grid inside the cage.
[0012] The installation of this additional part complicates the overall design of the assembly, in an already very dense environment. In addition, for the flexible cage to retain its deformation capacity, a radial clearance must be provided between the bars of the flexible cage and the annular grid. This clearance implies a possible passage of oil mist at the axial ends of the free annular space between the cage and the grid, with a risk of oil projection radially outwards against the hot wall of the stator part, passing through the large openings of the flexible cage.
[0013] Due to the oil projections observed on the hot part of the stator, and despite the presence of the annular grid supposed to limit these projections, the impacts cause forced convection type heat exchanges. This results in significant local increases in the oil temperature. These temperature increases harm the thermal balance of the oil, with a need to oversize the exchangers intended to cool it. In addition, causing excessively high local oil temperatures generates rapid thermal oxidation of the oil, leading to problems in taking into account the risks of coking.
[0014] STATEMENT OF THE INVENTION
[0015] To address at least partially the problems mentioned above relating to the embodiments of the prior art, the invention firstly relates to an assembly for an aircraft turbomachine, comprising a stator part, a second bearing, and a support part, the second bearing and the support part being arranged in an oil enclosure delimited radially outwards by an outer enclosure delimiting portion integrated into the stator part, the support part comprising:
[0016] - a first axial end portion preferably intended to form an outer ring of a first bearing, or to support such a ring;
[0017] - a fixing portion of the support part on the stator part, the fixing portion being connected to the first axial end portion;
[0018] - a second axial end portion opposite the first axial end portion, 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 part; and
[0019] - an intermediate ring, arranged axially between the first and second axial end portions, and forming a flexible connection between these same two axial end portions. Preferably, the intermediate ring, which is therefore arranged axially between the first and second axial end portions, is perforated so as to form a flexible connection between these same two axial end portions.
[0020] According to the invention, the intermediate ring further forms an oil splash protection element, configured to prevent oil from being splashed from inside this intermediate ring against the outer enclosure delimiting portion.
[0021] Thus, the intermediate ring can fulfill two functions, by means of this same part. The first function lies in 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 temperature of the oil, by preventing it from impacting the outer portion of the enclosure delimitation constituting a potentially hot zone of the turbomachine. Unlike the embodiments of the prior art in which two separate parts were required to ensure these two functions, the second of which was moreover not carried out optimally, the invention advantageously provides for merging these two parts into a single one, by perforating the ring much more finely in order to avoid unwanted projections of oil particles.
[0022] The design of the assembly is simplified, and the thermal function improved.
[0023] The invention furthermore has at least one of the following optional features, taken alone or in combination.
[0024] Preferably, the oil splash protection element is traversed by orifices, each made so that its smallest dimension is less than 5 mm.
[0025] According to a first preferred embodiment of the invention, the oil splash protection element has orifices succeeding one another in a circumferential direction of the ring, as well as in an axial direction thereof, these orifices being preferably arranged in annular rows of orifices succeeding one another axially. In this case, the splash protection element therefore takes the general shape of a grid. The orifices are preferably circular, oval or oblong in shape, or any other shape deemed appropriate.
[0026] According to a second preferred embodiment of the invention, the oil splash protection element has orifices succeeding 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.
[0027] Whatever the embodiment envisaged, the protective element has orifices as well as solid portions delimiting these orifices, and the ratio between the cumulative surface area of the orifices and the total surface area of the protective element is between 0.2 and 0.5. It may, for example, be of the order of 0.3 to 0.4.
[0028] Preferably, the attachment portion of the support piece to the stator portion takes the form of a attachment flange, or a plurality of attachment ears distributed circumferentially around the first axial end portion of the support piece.
[0029] Preferably, the outer enclosure delimiting portion, integrated into the stator part, has an oil recovery orifice in the lower part.
[0030] Preferably, the support part 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 part in several portions fixedly attached to each other, and / or by providing machining / drilling to make the orifices of the protective element.
[0031] The invention also relates to an aircraft turbomachine comprising at least one such assembly, preferably arranged downstream of a combustion chamber of the turbomachine.
[0032] 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.
[0033] Other advantages and characteristics of the invention will appear in the detailed non-limiting description below.
[0034] BRIEF DESCRIPTION OF THE DRAWINGS
[0035] This description will be made with regard to the attached drawings, including:
[0036] [Fig 1] represents a schematic axial sectional view of an aircraft turbomachine according to the invention;
[0037] [Fig 2] represents an enlarged schematic view in axial section 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;
[0038] [Fig 3] is an enlarged perspective view of the support piece equipping the assembly shown in the previous figure;
[0039] [Fig 4] is an axial sectional view of the support piece shown in the preceding figure;
[0040] [Fig 5] is a perspective view similar to that of Fig. 3, with the part being in the form of a second preferred embodiment of the invention; and [Fig 6] is an axial sectional view of the support part shown in the preceding figure.
[0041] DESCRIPTION OF EMBODIMENTS
[0042] Referring firstly to Figure 1, there is shown schematically a preferred embodiment of the invention, taking the form of an aircraft turbomachine 1, here a helicopter.
[0043] The turbomachine 1 comprises a main gearbox 2, intended to drive a rotating receiver (not shown), namely helicopter blades in this case. The main gearbox is connected to a gas generator 4 of the turbomachine, comprising, from upstream to downstream in a main flow direction 6 of the gases through the turbomachine, a centrifugal compressor 8, possibly double-stage, a combustion chamber 10, a high-pressure turbine 12, and a free turbine 14, also called a low-pressure turbine.
[0044] 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 gearbox 2, into which it enters to drive this gearbox. Subsequently, the shafts 16a, 16b will be called 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.
[0045] The invention relates more specifically to an assembly 20 located in the rear part of the turbomachine, being arranged in whole or in part downstream of the combustion chamber 10. The assembly 20 is preferably located between the two turbines 12, 14, and defines an oil enclosure 22 which is delimited radially towards the outside by an outer enclosure delimiting portion 24, this portion 24 belonging to a stator part 26 of the assembly 20. The stator part 26 is arranged axially between the two turbines, so that its outer enclosure delimiting portion 24 corresponds to an inter-turbine casing. The stator part 26, connected to the casing 25 of the turbines, takes the form of a housing with its outer annular portion 24 surrounding the shafts 16a, 16b.
[0046] In 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 first drive shaft 16a in rotation, while the second bearing 28b is a ball bearing guiding the second drive shaft 16b in rotation. In the invention, it is this second bearing 28b which is associated with a “squeeze film damper”, as will be detailed below.
[0047] As best seen in Figure 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 and 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 and 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.
[0048] Each of the two motor shafts 16a, 16b therefore penetrates axially into the stator part 24 and into the oil enclosure 22, the latter also containing a support part 32 specific to the invention, and a first embodiment of which is shown in Figures 1 to 4.
[0049] This support part 32 is preferably made in one piece rather than by assembling several components, even if the latter possibility remains conceivable. The part in one piece is made for example by any additive manufacturing technique, in a material preferably metallic, even if other conventional manufacturing techniques remain conceivable.
[0050] The support part 32 is generally annular and cylindrical, centered on the axis 18. This comprises the different elements described below:
[0051] - a first axial end portion 34a forming a support housing an outer ring 36 of the first bearing 28a, the inner ring of this bearing being carried externally by the first shaft 16a;
[0052] - a fixing portion 40 of the support part 32, on the outer annular portion 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 outer enclosure delimiting portion 24;
[0053] - a second axial end portion 34b opposite the first axial end portion 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 - 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.
[0054] The ring 46, which forms a flexible cage, is effectively finely perforated so as to be able to provide the flexibility required in the mechanical maintenance of the second axial end portion 34b. This second portion 34b, therefore 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.
[0055] The mechanical flexibility of the ring 46, combined with the damping oil film 50, effectively allows greater orbiting of the outer ring 34b in the stator portion 26 which houses it, and thus ensures better efficiency of the oil film compression damper. The latter is confined between two annular sealing tongues 51 spaced axially from each other, and carried radially inwards by the outer annular portion 24 of the stator portion 26.
[0056] Furthermore, it is noted that the outer annular portion 24 of the stator part 26 is also crossed by one or more oil injectors 53 in the enclosure, these injectors approaching for example as close as possible to the shafts 16a, 16b.
[0057] 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 integrating a second function of the oil thermal protection type. Indeed, the intermediate ring 46 also forms an oil splash protection element, configured to prevent oil from being sprayed from the enclosure from the inside of this intermediate ring, against the inner surface of the outer enclosure delimiting portion 24. In other words, the ring 46 has a certain porosity and is perforated sufficiently finely so that in operation, the particles of the 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 protection element 46.The oil projected against the latter can then obviously pass through this element, but its purpose in practice consists of reducing the speed of the oil or even stopping it completely in order to prevent it from impacting the outer portion of enclosure delimitation 24. After having passed through the protective element, the oil can nevertheless come into contact with a part of the inner surface of the outer portion of enclosure delimitation 24, for example after having trickled over the protective element and then having fallen by gravity onto or close to a lower part of the stator housing comprising one or more oil recovery orifices 54. The contact of the oil with the inner surface of the outer portion of enclosure delimitation 24 is therefore not made by projection of this oil from the inside of the ring 46, but it is above all partial because it only concerns a part of the aforementioned inner surface.Due to this partial contact, without oil projection onto the crankcase part, the increase in oil temperature by convection is advantageously contained. The risks of coking are thus reduced, and the oil cooling requirements are more measured, requiring less bulky exchangers.
[0058] In the first preferred embodiment of the invention, the intermediate ring 46 is finely perforated, presenting orifices 56 succeeding one another in a circumferential direction “C” of the ring, as well as in an axial direction “A” thereof. Thus, the orifices 56 are preferably arranged in annular rows of orifices succeeding one another axially, thus forming an anti-projection protection element in the form of a grid and matrix. A staggered arrangement is also conceivable. Here, the orifices 56 of the protection element 46 are circular in shape, but other shapes can be adopted, without departing from the scope of the invention. To ensure the desired anti-projection function, while offering the required flexibility, the orifices 56 are of small dimensions, and provided in high density.Also, the orifices 56 are each made so that its smallest dimension “Dmin”, that is to say 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 of the order of 2 mm. Furthermore, the density of orifices 56 within the element 46 can be of the order of four to five orifices / cm. 2 .
[0059] 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 follow one another 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 else having for example a circumferential component.
[0060] These orifices 56 are each made so that its smallest dimension “Dmin”, that is to say here the slot width in the circumferential direction “C”, is less than 5 mm. It can for example be of the order of 1.5 mm.
[0061] To provide the required flexibility while maintaining the small orifice width indicated above and necessary for the anti-oil splash function, the solid portions 58 in the form of bars which remain within the element 46 are preferably of small section. Moreover, the ratio between the cumulative surface area of the orifices 56 of this element, and the total surface area of the latter integrating the orifices 56 and the bars 58, is preferably between 0.2 and 0.5, being for example fixed at 0.4 in this preferred embodiment. This range of values concerning the proportion of the passage section of the protection element is also applicable to the first preferred embodiment described above, in which the ratio is for example fixed at 0.3. As an indicative example, the passage section offered by all of the orifices 56 is of the order of 2,000 mm 2. Of course, various modifications may be made by those skilled in the art to the invention which has just been described, solely by way of non-limiting examples and the scope of which 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
DEMANDS 1. Assembly (20) for an aircraft turbomachine, comprising a stator portion (26), a second bearing (28b), and a support piece (32), the second bearing (28b) and the support piece being arranged in an oil enclosure (22) radially delimited outwards by an outer portion of enclosure delimitation (24) integrated into the stator portion (26), the support piece (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 piece 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 portion (26); and - an intermediate ring (46), arranged axially between the first and second axial end portions (34a, 34b), and forming a flexible link between these two same axial end portions, characterized in that the intermediate ring (46) also forms an oil splash protection element, configured to prevent oil from being splashed from inside this intermediate ring (46), against the outer enclosure boundary portion (24).
2. Assembly according to claim 1, characterized in that the oil splash protection element (46) is traversed 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) following one another in a circumferential direction (C) of the ring, as well as in an axial direction (A) of the latter, these orifices (56) being preferably arranged in annular rows of orifices following one another axially.
4. Assembly according to claim 3, characterized in that the openings (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) following 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 area of the orifices and the total 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 lugs 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 portion of enclosure delimitation (24), integrated into the stator part (26), has in its lower part an oil recovery orifice (54).
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), and a second drive shaft (16b) guided by the second bearing (28b), the first and second shafts (28a, 28b) being concentric.