• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to a guide vane arranged in an air flow of a blower (15) of a turbofan aircraft engine, the aerodynamic part of the blade having an internal passage (50a) lubricant cooling part delimited by an intrados wall (70) and an extrados wall (72) of the blade. The passage is equipped with a thermal conduction matrix (80) compressed between the walls (70, 72) and separating a first lubricant circulation space (81a) from a second lubricant circulation space (81b). In addition, the matrix (80) defines on the one hand first contact members of the intrados wall (70) arranged in the first space (81a) and between which the lubricant of the first space is intended to circulate, and on the other hand, second contact members of the extrados wall (72) arranged in the second space (81b) and between which the lubricant of the second space is intended to circulate.
    公开号:FR3071008A1
    申请号:FR1758371
    申请日:2017-09-11
    公开日:2019-03-15
    发明作者:Cedric ZACCARDI；Christophe Marcel Lucien Perdrigeon；Mohamed-Lamine BOUTALEB；Dimitri Daniel Gabriel Marquie
    申请人:Safran Aircraft Engines SAS；
    IPC主号:
    专利说明:

    OUTPUT DIRECTIVE VANE FOR TURBOMACHINE, INCLUDING A LUBRICANT COOLING PASS EQUIPPED WITH A COMPRESSED THERMAL CONDUCTION MATRIX BETWEEN THE WALLS OF INTRADOS AND EXTRADOS
    DESCRIPTION
    TECHNICAL AREA
    The present invention relates to the field of aircraft turbomachines with double flow, and in particular to the design of guide vanes arranged in all or part of an air flow of a fan of the turbomachine.
    They are preferably outlet guide vanes, also called OGV (from the English “Outlet Guide Vane”), intended to straighten the air flow at the outlet of the blower. Alternatively or simultaneously, guide vanes could if necessary be placed at the inlet of the blower. The guide vanes are conventionally arranged in the secondary stream of the turbomachine.
    The invention preferably relates to an aircraft turbojet engine equipped with such outlet guide vanes.
    STATE OF THE PRIOR ART
    On certain double-flow turbomachines, it is known to install outlet guide vanes downstream of the fan to straighten the flow which escapes therefrom, and also possibly to fulfill a structural function. This latter function is in fact intended to allow the passage of the forces from the center of the turbomachine towards an outer shroud situated in the extension of the fan casing. In this case, an engine attachment is conventionally arranged on or near this outer shell, to ensure the attachment between the turbomachine and an aircraft pylon.
    Recently, it has also been proposed to assign an additional function to the output guide vanes. It is a heat exchanger function between the outside air passing through the crown of outlet guide vanes, and the lubricant circulating inside these vanes. This heat exchanger function is for example known from document US 8,616,834, or from document FR 2,989,110.
    The lubricant intended to be cooled by the outlet guide vanes can come from different areas of the turbomachine. It may indeed be a lubricant circulating through the lubrication chambers of the rolling bearings supporting the motor shafts and / or the fan hub, or else a lubricant dedicated to the lubrication of the mechanical transmission elements of the accessories box (from the English AGB "Accessory Geared Box"). Finally, it can also be used for the lubrication of a fan drive reduction gear, when such a reduction gear is provided on the turbomachine in order to reduce the speed of rotation of its fan.
    The growing needs for lubricant require adapting the heat dissipation capacity, associated with the exchangers intended for cooling the lubricant. The fact of assigning a role of heat exchanger to the outlet guide vanes, as in the solutions of the two documents cited above, in particular makes it possible to reduce, or even eliminate conventional exchangers of the ACOC type (from the English “ Air Cooled Oil Cooler ”), These ACOC exchangers being generally arranged in the secondary stream, their reduction / elimination makes it possible to limit the disturbances of the secondary flow, and thus to increase the overall efficiency of the turbomachine.
    Within the interior lubricant cooling passage, it is possible to install studs intended to disturb the flow of lubricant and to increase the wetted surface, in order to ensure better heat exchange. These studs are designed to be made in one piece with the body of the blade. Their distal end is covered by the closure cover, which is fixed to the body of the blade after these studs have been produced. In order to guarantee contact between the distal end of the studs and the closure cover, the manufacturing tolerances of the parts must be very precise, in particular as regards the height of the studs. This results in complexity of construction as well as significant manufacturing costs.
    Consequently, there is a need to arrive at a design which facilitates the manufacture of such a blade with an integrated exchanger.
    STATEMENT OF THE INVENTION
    To respond at least partially to this need, the invention firstly relates to a guide vane intended to be arranged in all or part of an air flow from a fan of an aircraft turbomachine with double flow, the guide vane comprising a foot, a head, as well as an aerodynamic flow straightening part arranged between the foot and the head of the vane, said aerodynamic part of the vane comprising at least one internal lubricant cooling passage partly delimited by a lower surface and by an upper surface of the vane, the lower surface forming part of a body of the vane and the upper surface forming part of a closing cover of this body, or vice versa.
    According to the invention, said interior passage is equipped with at least one thermal conduction matrix compressed between the lower and upper surfaces, said matrix separating on either side thereof a first circulation space for lubricant also delimited by the lower surface, of a second lubricant circulation space also delimited by the upper surface, and said matrix defining on the one hand first contact members of the lower surface arranged in the first space and between which the lubricant of the first space is intended to circulate, and on the other hand of the second contact members of the upper surface wall arranged in the second space and between which the lubricant of the second space is intended to circulate.
    The invention thus cleverly provides for bringing a thermal conduction matrix into the interior lubricant cooling passage. The material and the geometry of the matrix allow it to be compressed between the lower and upper surfaces, namely to be deformed compared to its initial shape adopted before its implantation in the dawn. This ensures contact between the matrix and the lower and upper surfaces, in order to guarantee better overall thermal performance without requiring high manufacturing constraints. Indeed, the deformation undergone by the matrix inside the lubricant circulation passage allows it to adapt to any irregularity between the lower and upper surfaces, while ensuring contact with them. Of course, the compression deformation desired for the matrix is made possible by a certain flexibility and / or deformability thereof.
    The invention therefore allows easier manufacture, reduced cost and increased heat exchange performance.
    The invention preferably provides at least any one of the following optional characteristics, taken individually or in combination.
    Said interior passage is equipped with several matrices of thermal conduction compressed between the intrados and extrados walls, said matrices being spaced from one another in a large direction so as to define, between any two of them and directly consecutive, a lubricant zone serving on the one hand for the collection of the lubricant coming from the first and second circulation spaces of the uppermost matrix, and on the other hand for the distribution of the lubricant towards the first and second circulation spaces of the most downstream matrix.
    In other words, in this preferred embodiment of the invention, there are provided discontinuities between the dies along the span direction of the blade, direction in which the lubricant circulates. This provides several benefits, the main ones of which are mentioned below.
    First of all, in the event of a significant difference between the flow passing through the first space and that passing through the second space of the same matrix, the introduction of one or more discontinuities makes it possible to rebalance the distribution of these flows during circulation of the lubricant in the interior passage of the blade.
    In addition, in the event of differential cooling between the lower and upper surfaces, the temperature of the lubricant in the two spaces results in a pressure differential between these two spaces. The direct consequence lies in the imbalance of the flow distribution, which is likely to degrade the overall heat exchange performance. The presence of one or more discontinuities in the matrices greatly reduces this risk.
    Also, the blade being twisted, its profile changes throughout the interior passage of the lubricant cooling. The thicknesses of passage of the fluid can therefore be caused to evolve along the span direction of the blade, and modify the pressure losses with the possible consequence of an imbalance in the distribution of the flow rates. Here again, the discontinuities mentioned above make it possible to reduce this risk.
    In conclusion on this preferred aspect of the invention, the introduction of discontinuities ensures better control of the distribution of flow rates between the first and second circulation spaces defined by the matrices, and consequently results in an increase in overall exchange performance. thermal.
    Preferably, the first and second contact members each have a generally frustoconical shape, with a narrowing section going towards its associated lower or upper surface.
    The first and second contact members each have a substantially planar contact end pressing against its associated lower or upper wall.
    It is possible to make the matrix more flexible at the place which will be crushed, that is to say at the top of the waves for the contact members each having a generally frustoconical shape. The matrix would then deform preferentially in these zones, rather than in the frustoconical sides.
    The invention also relates to an aircraft turbomachine, preferably a turbojet engine, comprising a plurality of guide vanes such as that described above, arranged downstream or upstream of a fan of the turbomachine.
    It also presents as an object a process for manufacturing such a blade, comprising the following steps:
    a) making said at least one thermal conduction matrix;
    b) placing the matrix in a part of said interior passage defined by the body of the blade;
    c) placing said closure cover on the blade body, so as to compress and deform the heat conduction matrix; and
    d) fixing of the closing cover on the blade body.
    Preferably, step a) is carried out by forming a sheet, preferably by cold forming. Alternatively, additive manufacturing could be envisaged, this type of production also being called 3D printing or direct manufacturing. The additive manufacturing of the heat conduction matrix can for example be carried out by any of the following techniques:
    - selective fusion by laser (from English “Selective Laser Melting” or “SLM”) or by electron beam (from English “Electron Beam Melting” or “EBM”);
    - selective sintering by laser (from “Selective Laser Sintering” or “SLS”) or by electron beam;
    - any other type of powder solidification technique under the action of a medium to high power source of energy, the principle being to melt or sinter a bed of metal powder by laser beam or electron beam.
    Preferably, step b) of the method is carried out so that the first and second contact members of the matrix each have a convex contact end, and so that after step c) of placing said cover closing, this contact end is substantially flat and bears against its associated pressure or pressure surface. There is thus a local deformation of these ends, favorable for obtaining certain and large contact surfaces.
    Step c) is preferably implemented so that the deformation of the matrix is an elastic deformation.
    Finally, step d) is preferably implemented by welding, brazing or bonding.
    Other advantages and characteristics of the invention will appear in the detailed non-limiting description below.
    BRIEF DESCRIPTION OF THE DRAWINGS
    This description will be made with reference to the accompanying drawings, among which;
    - Figure 1 shows a schematic side view of a turbojet engine according to the invention;
    - Figure 2 shows an enlarged view, more detailed, of an outlet guide vane of the turbojet engine shown in the previous figure, according to a first preferred embodiment of the invention;
    - Figure 3 is an enlarged perspective view of part of the outlet guide vane shown in the previous figure;
    - Figure 4 is a perspective view similar to that of the previous figure, from another angle of view and with the blade shown without its closure cover;
    - Figure 5 is a front view of part of the blade shown in the previous figure;
    - Figure 6 corresponds to a sectional view taken along the line VI-VI of Figure 5;
    - Figure 7 is a view similar to that of Figure 2, with the blade appearing according to a second preferred embodiment of the invention; and
    - Figures 8a to 8c illustrate different stages of a manufacturing process representative of the assembly of the blade shown in the previous figure.
    DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS
    Referring to Figure 1, there is shown a turbofan 1 with double flow and double body, having a high dilution rate. The turbojet engine 1 conventionally comprises a gas generator 2 on either side of which are arranged a low pressure compressor 4 and a low pressure turbine 12, this gas generator 2 comprising a high pressure compressor 6, a combustion chamber 8 and a high pressure turbine 10. Subsequently, the terms “front” and “rear” are considered in a direction 14 opposite to the main direction of flow of the gases within the turbojet engine, this direction 14 being parallel to the axis. longitudinal 3 thereof. On the other hand, the terms “upstream” and “downstream” are considered according to the main direction of flow of the gases within the turbojet engine.
    The low pressure compressor 4 and the low pressure turbine 12 form a low pressure body, and are connected to each other by a low pressure shaft 11 centered on the axis 3. Likewise, the high pressure compressor 6 and the high pressure turbine 10 form a high pressure body, and are connected to each other by a high pressure shaft 13 centered on the axis 3 and arranged around the low pressure shaft 11. The shafts are supported by bearings bearing 19, which are lubricated by being arranged in oil chambers. It is the same for the fan hub 17, also supported by rolling bearings 19.
    The turbojet engine 1 also comprises, at the front of the gas generator 2 and of the low pressure compressor 4, a single blower 15 which is here arranged directly behind a cone of air intake of the engine. The fan 15 is rotatable along the axis 3, and surrounded by a fan casing 9. In FIG. 1, it is not driven directly by the low pressure shaft 11, but only indirectly driven by this shaft via a reducer 20, which allows it to rotate with a slower speed. Nevertheless, a solution with direct drive of the fan 15, by the low pressure shaft 11, comes within the scope of the invention.
    In addition, the turbojet engine 1 defines a primary stream 16 intended to be traversed by a primary flow, as well as a secondary stream 18 intended to be crossed by a secondary stream located radially outward relative to the primary stream, the stream of the fan being therefore divided. As is known to a person skilled in the art, the secondary duct 18 is delimited radially outwards in part by an external ferrule 23, preferably metallic, extending rearward the fan casing 9.
    Although this has not been shown, the turbojet engine 1 is equipped with a set of equipment, for example of the fuel pump, hydraulic pump, alternator, starter, variable-timing stator (VSV) actuator, valve actuator or electric power generator. This is in particular an equipment for the lubrication of the reduction gear 20. This equipment is driven by an accessories box or AGB (not shown), which is also lubricated.
    Downstream of the fan 15, in the secondary vein 18, there is provided a crown of guide vanes which are here outlet guide vanes 24 (or OGV, from the English “Outlet Guide Vane”). These stator vanes 24 connect the outer shell 23 to a casing 26 surrounding the low pressure compressor 4. They are spaced circumferentially from one another, and make it possible to straighten the secondary flow after it has passed through the blower 15. In addition, these vanes 24 can also fulfill a structural function, as is the case in exemplary embodiments which are presently described. They transfer the forces coming from the reduction gear and the rolling bearings 19 of the motor shafts and of the fan hub, to the outer shell 23. Then, these forces can pass through a motor attachment 30 fixed on the shell 23 and connecting the turbojet engine. to an attachment pylon (not shown) of the aircraft.
    Finally, the outlet guide vanes 24 provide, in the embodiments which are presently described, a third function of heat exchanger between the secondary air flow passing through the crown of blades, and of the lubricant circulating inside these vanes 24. The lubricant intended to be cooled by the outlet guide vanes 24 is that used for the lubrication of the rolling bearings 19, and / or the equipment of the turbojet engine, and / or the accessories box, and / or the reducer 20. These vanes 24 thus form part of the fluid circuit (s) in which the lubricant is circulated in order to successively lubricate the associated element (s), then to be cooled.
    With reference now to FIGS. 2 to 6, one of the outlet guide vanes 24 will be described, according to a first preferred embodiment of the invention. In this regard, it is noted that the invention as it will be described with reference to Figures 2 to 6 can be applied to all the vanes 24 of the stator ring centered on the axis 3, or only to certain of these blades.
    The blade 24 can be of strictly radial orientation as in FIG. 1, or else be slightly inclined axially as shown in the FIG.
    2. In all cases, it is preferably straight in side view as shown in FIG. 2, extending in a direction of span 25.
    The outlet guide vane 24 has an aerodynamic part 32 which corresponds to its central part, that is to say that exposed to the secondary flow. On either side of this aerodynamic part 32 serving to straighten the flow leaving the fan, the blade 24 comprises a foot 34 and a head 36 respectively.
    The foot 34 is used for fixing the blade 24 on the housing of the low pressure compressor, while the head is used for fixing the same blade on the outer shell extending the fan housing. In addition, the blade 24 comprises at the level of its foot and of its head, platforms 40 serving to reconstitute the secondary vein between the blades 24, in the circumferential direction.
    The aerodynamic part 32 of the blade is preferably manufactured in two separate parts, then fixedly attached to one another. It is first of all a body of the blade 32a, which includes not only a large part of the aerodynamic part 32, but also the foot 34, the head 36 and the platforms 40. This body 32a is produced in one piece. The other part is formed by a cover 32b closing the body, and fixed to the latter by a conventional technique such as welding, soldering or bonding.
    In this first preferred embodiment of the invention, the aerodynamic part 32 is equipped with two interior passages 50a, 50b substantially parallel to each other, and parallel to the span direction 25. More precisely, it s 'Acts of a first interior passage 50a of lubricant cooling, which extends in a first main direction 52a of lubricant flow. This direction 52a is substantially parallel to the span direction 25, and has a direction going from the foot 34 towards the head 36. In a similar manner, a second interior passage 50b for cooling the lubricant is provided, which extends in a second main direction 52b of lubricant flow within this passage. This direction 52b is also substantially parallel to the span direction 25, and has an opposite direction going from the head 36 to the foot 34. The first passage 50a is therefore intended to be crossed radially outward by the lubricant, while the second passage 50b is intended to be crossed radially inward. To ensure the passage from one to the other, near the head 36, the external radial ends of the two passages 50a, 50b are fluidly connected by an elbow 54 at 180 °, corresponding to a hollow formed in the aerodynamic part 32. Alternatively, the passages 50a, 50b do not connect within the aerodynamic part 32 of the blade 24, but each extend separately over the entire length of the aerodynamic part 32. To fluidly connect one to the other outside of the blade 24, there is for example provided a connection elbow arranged radially outward relative to the blade head 36, for example resting on this head.
    The internal radial ends of the two passages 50a, 50b are in turn connected to the lubricant circuit 56, shown diagrammatically by the element 56 in FIG. 2. This circuit 56 notably comprises a pump (not shown), making it possible to apply to the lubricant the desired direction of circulation within the passages 50a, 50b, namely the introduction of the lubricant by the internal radial end of the first passage 50a, and the extraction of the lubricant by the internal radial end of the second passage 50b. Connections 66 ensure fluid communication between the internal radial ends of the passages 50a, 50b and the circuit 56, these connections 66 passing through the foot 34.
    The two passages 50a, 50b and the elbow 54 together have a general U shape, with the first passage 50a and the second passage 50b offset from one another in a transverse direction 60 of the blade substantially orthogonal to the large-scale direction 25. To optimize the best heat exchange, the first passage 50a is located on the side of a trailing edge 62 of the blade 24, while the second passage 50b is located on the side of an edge d Attack 64. However, a reverse situation can be retained, without departing from the scope of the invention. It is also noted that the invention could provide an aerodynamic part 32 with only a single internal cooling passage, without departing from the scope of the invention. In this case, certain blades would be crossed by the lubricant from the inside to the outside, while other blades would be crossed in the opposite direction.
    The aerodynamic part 32 of the outlet guide vane 24 comprises a lower surface 70, an upper surface 72, a solid area 74 connecting the two walls 70, 72 near the trailing edge 62, a solid area 76 connecting the two walls 70, 72 near the leading edge 64, as well as a central solid area 78. This latter area 78 connects the two walls 70, 72 at a substantially central portion thereof, according to the direction of the dawn rope. It also serves as structural reinforcement and extends from the foot 34 to the elbow 54, while the solid areas 74, 76 extend over substantially the entire length of the part 32, in the span direction 25. The first passage 50a is formed between the walls 70, 72 and between the solid areas 74, 78, while the second passage 50b is formed between the walls 70, 72 and between the solid areas 76, 78. The walls of the underside and of upper surfaces 70, 72 have, with regard to the passages 50a, 50b which they delimit, substantially constant thicknesses. On the other hand, the passages 50a, 50b extend transversely in the direction 60 with a variable thickness between the two walls 70, 72. The maximum thickness of these passages can be of the order of a few millimeters. Alternatively, the passages 50a, 50b could have a constant thickness, but in this case the two walls 70, 72 would then adopt a variable thickness to obtain the aerodynamic profile of the blade.
    It is noted that the upper surface 72 is integrated into the body 32a of the blade, while the lower surface 70 is integrated into the cover 32b, the latter extending between the solid areas 74, 76 which it forms in part. In this regard, the solid zones 74, 76 have recesses 77 forming support and fixing zones for the cover 32b on the body 32a. These recesses 77, of depth substantially equal to the thickness of the closure cover 32b, allow a flush aerodynamic junction between these two components 32a, 32b.
    The two interior passages 50a, 50b for cooling the lubricant have the particularity of integrating one or more thermal conduction matrices 80. The presence of matrices makes it possible to improve the heat exchange performance, in particular thanks to the fact that it provides an increase in the wetted surface on the side of the lubricant which passes through the passages 50a, 50b. This matrix 80 also makes it possible to disturb the passage of the lubricant, thereby generating turbulence which directly influences the convection coefficient of the lubricant passing through the matrix. The definition of such a matrix can thus be carried out so as to maximize the exchange performance while creating the least pressure losses possible between the entry and the exit of the blade.
    To improve thermal performance, the matrix 80 can be made of a material different from that of the body 32a and the cover 32b. As indicative examples, the matrix 80 can be produced using an alloy based on aluminum or titanium. Its overall volume may have a thickness of the order of several millimeters, which corresponds to the thickness of the interior passage in which it is housed.
    In the first preferred embodiment, a single thermal conduction matrix 80 is provided in each interior passage 50a, 50b. These two dies 80 are of substantially identical or similar designs. They also have identical or similar densities of contact members, even if it could be otherwise, without departing from the scope of the invention. Consequently, only the matrix 80 of the first interior passage 50a will now be described, but it should be understood that this description is also applicable by analogy to the matrix of the second interior passage 50b. Furthermore, it is noted that the elbow 54 defines an interior space which is preferably free of studs. It remains empty, or can be equipped with concentric walls channeling the lubricant from one passage to the other.
    The matrix 80 has the particularity of being compressed between the lower surface 70 and upper surface 72. As a result, the contact of the matrix 80 with these two walls is ensured, and the thermal performance is increased. The matrix 80 takes the form of a sheet of complex shape which separates a first lubricant circulation space 81a from a second lubricant circulation space 81b. The first space 81a is also delimited by the lower surface 70 and by the solid areas 74, 78. The second space 81b is also delimited by the upper surface 72 and by the solid zones 74, 78. Consequently, the lubricant which enters the first space 81a cannot join the second space 81b before it leaves the matrix, and vice versa.
    The matrix 80 first of all defines first contact members of the lower surface, referenced 82a. These members 82a are substantially orthogonal to the direction 52a. They each have a generally frustoconical shape with an axis substantially orthogonal locally to the lower surface wall 70, and with a section which tapers towards this wall. Each member 82a ends in a substantially planar contact end 84a, bearing against the pressure surface 70. This substantially planar end corresponds to a part of the matrix which is elastically deformed due to the compression of the latter between the body 32a and its closing cover 32b. Its diameter is of the order of 0.5 to 10 mm.
    Similarly, the matrix 80 then defines second contact members of the lower surface, referenced 82b. These members 82b are substantially orthogonal to direction 52a. They each have a generally frustoconical shape with an axis substantially orthogonal locally to the upper surface wall 72, and with a section which tapers towards this wall. Each member 82b ends in a substantially planar contact end 84b, pressing against the upper surface 72. This substantially planar end corresponds to a part of the matrix which is elastically deformed due to the compression of the latter between the body 32a and its closing cover 32b. Its diameter is also of the order of 0.5 to 10 mm.
    The first contact members 82a are located in the first lubricant circulation space 81a, within which the lubricant is intended to flow between these members 82a arranged in rows. In the same way, the second contact members 82b are located in the second lubricant circulation space 81b, within which the lubricant is intended to flow between these second members 82b also arranged in rows. In the span direction 25, the rows of first members 82a are arranged alternately with the rows of second members 82b, preferably so that all of the contact ends 84a, 84b are arranged in staggered rows, like this is best seen in Figures 4 and 5.
    In at least one zone of the passage 50a, and preferably in the entirety of the latter, the contact members 82a, 82b form a set of studs provided in a density for example of around 3 studs / cm 2 . More generally, the density is for example between about 1 and 5 pads / cm 2 on average.
    Returning to FIG. 2, during the operation of the engine, the lubricant circulating through the circuit 56 is introduced into the first interior passage 50a, in the first direction 52a going radially outward. At this point, the lubricant has a high temperature. A heat exchange then takes place between this lubricant conforming to the matrix 80 of the first passage 50a, and the secondary flow conforming to the external surface of the pressure and pressure surfaces. The lubricant, after having been redirected by the elbow 54 in the second passage 50b, undergoes in the latter a similar cooling, always by heat exchange with the secondary air flow and by circulating in the second main direction of flow 52b. Then, the cooled lubricant is extracted from the blade 24, and redirected by the closed circuit 56 towards elements to be lubricated and / or towards a reservoir of lubricant from which cooled lubricant is pumped to lubricate elements.
    FIG. 7 shows a second preferred embodiment in which each interior passage 50a, 50b is equipped with several matrices of thermal conduction compressed between the walls of lower surface 70 and upper surfaces 72.
    In each passage 50a, 50b, the dies 80 follow one another in the span direction 25, corresponding to the main directions of flow of the lubricant 52a, 52b. These dies 80 are spaced from each other in the span direction 25 so as to define, between any two of them and directly consecutive, a lubricant area 86 whose section corresponds to that of the associated passage 50a, 50b. In other words, each lubricant zone 86 is preferably left free to serve for collecting the lubricant coming from the first and second circulation spaces of the matrix 80 most upstream in the direction of flow of the lubricant. This zone 86 then also serves for the distribution of the lubricant to the first and second circulation spaces of the matrix 80 directly consecutive, downstream. Consequently, these zones of lubricant 86 form discontinuities in the heat exchange structure, which in particular makes it possible to rebalance the distribution of the flow rates between the first and second circulation spaces, before the lubricant enters each new matrix 80.
    FIGS. 8a to 8c show different stages of a process for manufacturing a blade according to any one of the embodiments presented above.
    First of all, there is provided a step for producing each die 80, preferably by cold forming a sheet 80 ′ initially planar. This sheet 80 ', shown in Figure 8a, has a thickness of about 0.08 to 3 mm. Its forming implies for example a deformation of the order of 50 to 70%, so as to reveal the contact members 82a, 82b. At the end of this forming, the contact ends 84a, 84b are not flat, but curved.
    During a next step, called step b), the dies 80 are placed in a part 88 of the interior passage defined by the body of the blade 32. This step b) is shown diagrammatically in FIG. 8b . Once housed in part 88 of the passage, the die 80 in the non-constrained state extends beyond the delimitation intended to be formed subsequently by the closure cover 32b, this delimitation being shown diagrammatically by the dotted line 90. By way of example, the matrix 80 can exceed the delimitation 80 by a height H which can range up to 5 mm.
    Then, a step c) is carried out of placing the closing cover 32b on the blade body 32, by applying a force 92 high enough on the cover 32b so that the latter comes to bear against the recesses 77 of the body 32a. This force 92 is, for example, of the order of 10 to 10,000 N. During the fitting of the cover 32b, each matrix 80 compresses and elastically deforms between the walls of lower surface 70 and upper surfaces 72. During this phase, the contact ends 84a, 84b deform locally, gradually losing their domed shape to adopt a substantially planar shape, matching the interior surfaces of the walls 70, 72 over a larger area. This step c), shown diagrammatically in FIG. 8c, is possible thanks to the relative flexibility of the matrix 80, which is therefore capable of deforming elastically and locally at the contact ends, in order to adapt perfectly to the distance separating the two walls 70, 72. It is noted that part of the deformation can also be plastic, without departing from the scope of the invention.
    Finally, step d) consists in fixing the closure cover to the blade body, preferably by gluing, welding or brazing.
    Of course, various modifications can be made by those skilled in the art to the invention which has just been described, only by way of nonlimiting examples. In particular, the technical characteristics specific to each of the embodiments described above can be combined with one another, without departing from the scope of the invention. Finally, it is noted that in the non-illustrated case of the inlet guide vanes to straighten the air flow upstream of the blower, these blades are arranged throughout the air flow of the blower around a cone d non-rotating air inlet, the feet of the blades then being connected to this fixed air inlet cone.
    Furthermore, other engine architectures also fall within the scope of the invention, by responding to the designation "turbomachine of an aircraft with double flow". It can for example be a triple body architecture by (namely comprising three shafts respectively connecting first stages of turbine to a blower, second stages of turbine to stages of low pressure compressor, and third stages of turbine with high pressure compressor stages), or even a non-faired blower architecture (called "open rotor") comprising rectifiers, such as a motor having an open "secondary flow" including a fast propeller fulfilling the role of blower as well as a fixed stator rectifier arranged at the rear, but without external casing.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    1. Directing vane (24) intended to be arranged in all or part of an air flow of a fan (15) of an aircraft turbomachine with double flow, the directing vane comprising a foot (34), a head (36), as well as an aerodynamic flow straightening part (32) arranged between the foot and the head of the blade, said aerodynamic part of the blade comprising at least one internal cooling passage (50a, 50b) lubricant partly delimited by a lower surface wall (70) and by a lower surface wall (72) of the blade, the lower surface wall (70) forming part of a body (32a) of the blade and the upper surface wall (72) forming part of a closure cover (32b) of this body, or vice versa, characterized in that said internal passage (50a, 50b) is equipped with at least one thermal conduction matrix (80) compressed between the lower surface (70) and upper surface (72), said matrix separating on either side thereof a first lubrication circulation space ant (81a) also delimited by the lower surface (70), of a second lubricant circulation space (81b) also delimited by the upper surface (72), and in that said matrix (80) defines on the one hand of the first members (82a) of contact with the pressure surface (70) arranged in the first space (81a) and between which the lubricant of the first space is intended to circulate, and on the other hand of the second members (82b) for contacting the upper surface wall (72) arranged in the second space (81b) and between which the lubricant of the second space is intended to circulate.
    [2" id="c-fr-0002]
    2. Dawn according to claim 1, characterized in that said interior passage (50a, 50b) is equipped with several thermal conduction matrices (80) compressed between the walls of the lower surface (70) and upper surfaces (72), said dies being spaced from each other in a major direction (25) so as to define, between any two of them and directly consecutive, a lubricant zone (86) used on the one hand for collecting the lubricant from first and second circulation spaces (81a, 81b) of the matrix (80) most upstream, and on the other hand to the distribution of the lubricant towards the first and second circulation spaces (81a, 81b) of the matrix (80 ) the most downstream.
    [3" id="c-fr-0003]
    3. Dawn according to any one of the preceding claims, characterized in that the first and second contact members (82a, 82b) each have a generally frustoconical shape, with a narrowing section going towards its lower surface or d associated upper surface.
    [4" id="c-fr-0004]
    4. Dawn according to any one of the preceding claims, characterized in that the first and second contact members (82a, 82b) each have a substantially planar contact end (84a, 84b) bearing against its pressure surface or associated upper surface.
    [5" id="c-fr-0005]
    5. aircraft turbomachine (1), preferably a turbojet engine, comprising a plurality of guide vanes (24) according to any one of the preceding claims, arranged downstream or upstream of a fan (15) of the turbomachine .
    [6" id="c-fr-0006]
    6. A method of manufacturing a blade (24) according to any one of claims 1 to 4, characterized in that it comprises the following steps:
    a) making said at least one thermal conduction matrix (80);
    b) placing the matrix (80) in a part (88) of said interior passage (50a, 50b) defined by the body (32a) of the blade;
    c) placing said closure cover (32b) on the blade body (32a), so as to compress and deform the heat conduction matrix (80); and
    d) fixing the closing cover (32b) to the blade body (32a).
    [7" id="c-fr-0007]
    7. Method according to claim 6, characterized in that step a) is carried out by forming a sheet (80 '), preferably by cold forming.
    [8" id="c-fr-0008]
    8. Method according to claim 6 or claim 7, aimed at manufacturing the blade according to claim 4, characterized in that step b) is carried out so that the first and second contact members (82a, 82b) of the matrix (80) each have a curved contact end (84a, 84b), and so that after step c) of placing said closure cover, this contact end (84a, 84b) is substantially planar and resting against its associated lower or upper wall.
    [9" id="c-fr-0009]
    9. Method according to any one of claims 6 to 8, characterized in that step c) is implemented so that the deformation of the matrix (80) is an elastic deformation.
    [10" id="c-fr-0010]
    10. Method according to any one of claims 6 to 9, characterized in that step d) is implemented by welding, brazing or gluing.
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    同族专利:
    公开号 | 公开日
    US11015468B2|2021-05-25|
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    GB201814724D0|2018-10-24|
    GB2568143A|2019-05-08|
    US20190078452A1|2019-03-14|
    GB2568143B|2022-03-16|
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    FR3046811A1|2016-01-15|2017-07-21|Snecma|DAUGHTER OUTPUT DIRECTOR FOR AIRCRAFT TURBOMACHINE, HAVING AN IMPROVED LUBRICANT COOLING FUNCTION|
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    FR2978196B1|2011-07-20|2016-12-09|Snecma|TURBOMACHINE AUB COMPRISING A PLATE REPORTED ON A MAIN PART|
    FR2989110B1|2012-04-05|2016-09-09|Snecma|DAWN OF STATOR FORMED BY A SET OF DAWN PARTS|
    US10196932B2|2015-12-08|2019-02-05|General Electric Company|OGV heat exchangers networked in parallel and serial flow|
    FR3049644B1|2016-04-01|2018-04-13|Safran Aircraft Engines|AIRBORNE TURBOMACHINE EXIT OUTPUT AUBE, HAVING AN IMPROVED LUBRICANT COOLING FUNCTION USING A THERMAL CONDUCTION MATRIX OCCURRING IN AN INTERIOR PASSAGE OF THE DAWN|
    US11168583B2|2016-07-22|2021-11-09|General Electric Company|Systems and methods for cooling components within a gas turbine engine|
    FR3059353B1|2016-11-29|2019-05-17|Safran Aircraft Engines|AIRBOARD TURBOMACHINE EXIT OUTPUT AUDE COMPRISING A LUBRICANT-BENDED ZONE HAVING AN IMPROVED DESIGN|
    FR3063767B1|2017-03-13|2019-04-26|Safran Aircraft Engines|OUTPUT DIRECTOR FOR AIRCRAFT TURBOMACHINE WITH IMPROVED LUBRICANT COOLING FUNCTION|
    FR3071008B1|2017-09-11|2019-09-13|Safran Aircraft Engines|DRAFT OUTPUT DIRECTOR FOR TURBOMACHINE, COMPRISING A LUBRICANT COOLING PASSAGE EQUIPPED WITH COMPRESSED THERMAL CONDUCTION MATRIX BETWEEN THE INTRADOS AND EXTRADOS WALLS|
    FR3077850B1|2018-02-13|2020-03-13|Safran Aircraft Engines|OUTPUT DIRECTIVE VANE FOR TURBOMACHINE, MADE FROM SEVERAL PARTS ASSEMBLED BETWEEN THEM BY MEANS OF ATTACHMENT OF THE VEIN|DE102015110615A1|2015-07-01|2017-01-19|Rolls-Royce Deutschland Ltd & Co Kg|Guide vane of a gas turbine engine, in particular an aircraft engine|
    FR3059353B1|2016-11-29|2019-05-17|Safran Aircraft Engines|AIRBOARD TURBOMACHINE EXIT OUTPUT AUDE COMPRISING A LUBRICANT-BENDED ZONE HAVING AN IMPROVED DESIGN|
    FR3066532B1|2017-05-22|2019-07-12|Safran Aircraft Engines|AIRBOARD TURBOMACHINE EXIT OUTPUT AUBE, COMPRISING A LUBRICANT COOLING PASS WITH FLOW-MAKING FLUID DISRUPTORS OF SIMPLIFIED MANUFACTURING|
    FR3071008B1|2017-09-11|2019-09-13|Safran Aircraft Engines|DRAFT OUTPUT DIRECTOR FOR TURBOMACHINE, COMPRISING A LUBRICANT COOLING PASSAGE EQUIPPED WITH COMPRESSED THERMAL CONDUCTION MATRIX BETWEEN THE INTRADOS AND EXTRADOS WALLS|
    FR3075256B1|2017-12-19|2020-01-10|Safran Aircraft Engines|OUTPUT DIRECTIVE VANE FOR AIRCRAFT TURBOMACHINE, INCLUDING A LUBRICANT COOLING PASS EQUIPPED WITH FLOW DISTURBORING PADS|
    FR3077850B1|2018-02-13|2020-03-13|Safran Aircraft Engines|OUTPUT DIRECTIVE VANE FOR TURBOMACHINE, MADE FROM SEVERAL PARTS ASSEMBLED BETWEEN THEM BY MEANS OF ATTACHMENT OF THE VEIN|
    FR3081912B1|2018-05-29|2020-09-04|Safran Aircraft Engines|TURBOMACHINE VANE INCLUDING AN INTERNAL FLUID FLOW PASSAGE EQUIPPED WITH A PLURALITY OF DISTURBING ELEMENTS WITH OPTIMIZED LAYOUT|
    CN109973224A|2019-05-08|2019-07-05|中国航空发动机研究院|Aircraft engine oil cooling system and method|
    法律状态:
    2019-03-15| PLSC| Publication of the preliminary search report|Effective date: 20190315 |
    2019-08-20| PLFP| Fee payment|Year of fee payment: 3 |
    2020-08-19| PLFP| Fee payment|Year of fee payment: 4 |
    2021-08-19| PLFP| Fee payment|Year of fee payment: 5 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1758371A|FR3071008B1|2017-09-11|2017-09-11|DRAFT OUTPUT DIRECTOR FOR TURBOMACHINE, COMPRISING A LUBRICANT COOLING PASSAGE EQUIPPED WITH COMPRESSED THERMAL CONDUCTION MATRIX BETWEEN THE INTRADOS AND EXTRADOS WALLS|
    FR1758371|2017-09-11|FR1758371A| FR3071008B1|2017-09-11|2017-09-11|DRAFT OUTPUT DIRECTOR FOR TURBOMACHINE, COMPRISING A LUBRICANT COOLING PASSAGE EQUIPPED WITH COMPRESSED THERMAL CONDUCTION MATRIX BETWEEN THE INTRADOS AND EXTRADOS WALLS|
    US16/126,028| US11015468B2|2017-09-11|2018-09-10|Outlet guide vane for turbomachine, comprising a lubricant cooling passage equipped with a thermal conducting matrix compressed between the intrados and extrados walls|
    GB1814724.9A| GB2568143B|2017-09-11|2018-09-11|Outlet guide vane for turbomachine, comprising a lubricant cooling passage equipped with a thermal conducting matrix compressed between the intrados and|
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