• 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 for a dual-flow aircraft turbomachine, the aerodynamic portion (32) of which has an inner lubricant cooling passage (50a) extending along a main lubricant flow direction (52a). According to the invention, the aerodynamic portion is made in one piece and also comprises heat transfer fins (80a, 80b) arranged in the passage (50a) connecting the intrados and extrados walls (70, 72). ) and extending substantially parallel to the direction (52a), these fins being distributed in successive rows in the main direction (52a) and made so that for any two rows of fins (R1, R2) directly consecutive, the row (R1) comprises fins (80a) forming a positive acute angle A1 with a fictitious reference plane of the blade (Pf), while the row (R2) comprises fins (80b) forming a negative acute angle A2 with this plan (Pf).
    公开号:FR3063767A1
    申请号:FR1752017
    申请日:2017-03-13
    公开日:2018-09-14
    发明作者:Mohamed-Lamine BOUTALEB；Fabien Roger Gaston Caty；Sebastien Vincent Francois Dreano；Thierry Georges Paul Papin；Christophe Marcel Lucien Perdrigeon；Cedric ZACCARDI
    申请人:Safran Aircraft Engines SAS；
    IPC主号:
    专利说明:

    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 blower 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 shroud, to ensure 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, makes it possible in particular 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.
    However, the solutions proposed in the prior art remain perfectible. In particular, there is a need to improve heat exchange in order to further increase the heat dissipation capacity. There is also a need to strengthen the mechanical strength and the tightness, vis-à-vis the high pressures generated by the circulation of the lubricant within these blades. This need to strengthen the mechanical strength is even more important in the particular case of a guide vane with a structural function. Finally, 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 these needs, the subject of the invention is first of all a guide vane intended to be arranged in all or part of an air flow of a fan of an aircraft turbomachine with double flow, the guide blade comprising a foot, a head, as well as an aerodynamic flow straightening part arranged between the foot and the head of the blade, said aerodynamic part of the blade comprising a first internal lubricant cooling passage s extending in a first main direction of flow of the lubricant from the foot towards the blade head, said first interior passage being partly delimited by a lower surface wall and by an upper surface wall of the blade.
    According to the invention, the aerodynamic part of the blade is made in one piece and also includes heat transfer fins arranged in the first passage by connecting the walls of lower and upper surfaces and extending substantially parallel in the first direction, said fins being distributed in rows of fins succeeding each other in the first main direction and produced so that for a first and a second row of any fins directly consecutive, the first row comprises at least several fins forming a positive acute angle A1 with a fictitious reference plane of the blade parallel to the first direction, while the second row comprises at least several fins forming a negative acute angle A2 with said fictitious reference plane.
    Thanks to the manufacture in one piece, the invention overcomes the problems of sealing and mechanical strength encountered in the known embodiments of the prior art, in particular in the solution described in document US 8,616. 834. In addition, the inverted inclinations of the fins which connect the lower and upper surfaces provide improved mechanical strength by behaving like a trellis.
    The fact of making the fins in one piece with the lower and upper walls also promotes better thermal conduction in the direction of these walls. In this regard, it is added that the various inclinations of the fins as well as their arrangement in rows provide high thermal performance, while limiting the pressure losses undergone by the lubricant passing through the first interior passage equipped with the fins.
    Finally, it is noted that thanks to the inclination of the fins according to the angles A1 and A2, these can be produced easily by additive manufacturing, in particular by orienting the blade to be manufactured so that its fictitious reference plane is parallel to the support surface of the blade during this manufacture.
    The invention also has at least one of the following optional features, taken individually or in combination.
    All the blades of the blade each form a positive acute angle A1 or a negative acute angle A2 with the fictitious reference plane, these angles A1, A2 being between 30 and 60 °.
    According to a first preferred embodiment of the invention, for a first and a second row of any directly consecutive fins, all the fins of the first row each form a positive acute angle A1 with the fictitious reference plane, while all the fins of the second row each form an acute negative angle A2 with said fictitious reference plane. For information, it is noted that in this description, the positive direction here corresponds to the clockwise from the fictitious reference plane, while the negative direction corresponds to the counterclockwise direction. However, an opposite solution could be adopted, without departing from the scope of the invention.
    Preferably, the positive acute angle A1 is substantially identical for all the fins of the first row, while the negative acute angle A2 is substantially identical for all the fins of the second row. However, the value of the angle could vary within each row, without departing from the scope of the invention.
    In addition, it is preferably made so that the fins of the first row are regularly spaced from one another in a transverse direction of the blade going from a leading edge towards a trailing edge of its aerodynamic part, that the fins of the second row are regularly spaced from each other in the transverse direction, and that in view in the first direction, the fins of the first row are arranged between the fins of the second row, preferably to jointly form a broken line. However, still in view in this first direction, gaps and / or intersections between the fins can be adopted, without departing from the scope of the invention.
    According to a second preferred embodiment of the invention, for a first and a second row of any directly consecutive fins, each of these rows comprises alternately, in a transverse direction of the blade going from a leading edge towards a trailing edge of its aerodynamic part, fins each forming a positive acute angle A1 with the fictitious reference plane and fins each forming a negative acute angle A2 with said fictitious reference plane.
    Preferably, in a view in the first direction, the fins of the first row jointly form a first broken line and the fins of the second row jointly form a second broken line, the first and the second broken line being offset one of the 'other in the transverse direction so that at least some fins of the first row cross at least some fins of the second row.
    Even more preferably, the first and the second broken line are periodic with the same period T, and they are offset from one another in the transverse direction by a value of T / n, n being a positive integer greater than 1, preferably between two and four.
    Whatever the embodiment envisaged, the first interior passage preferably comprises at least one zone in which said heat transfer fins are provided in a density of between 1 and 5 fins / cm 2 . This density can be uniform in the first pass, or variable. This ability to vary the density of fins makes it possible in particular to locally adapt the heat exchanges between the lubricant and the secondary flow.
    The dawn could comprise only a single first passage, ensuring the circulation of the lubricant radially towards the outside. In this case, the crown of guide vanes would then comprise at least one other vane of similar design, with an internal passage ensuring the circulation of the lubricant radially inwards.
    However, the aerodynamic part of the blade preferably also includes a second internal lubricant cooling passage extending in a second main direction of flow of the lubricant going from the head to the foot of the blade. This second interior passage is then preferably equipped with fins arranged according to a principle identical or similar to that observed in the first interior passage.
    According to one possibility, the first and second interior passages each extend separately through the entire aerodynamic part of the blade.
    According to another possibility, the first and second interior passages are fluidly connected to each other near the head of the blade, and the average density of fins within the first interior passage is then preferably less than the density of fins within the second interior passage. Indeed, since the lubricant is cooler in the direction of return adopted within the second interior passage, it is thus possible to increase the heat power exchanged by increasing the average density of the fin in this second passage.
    Preferably, the guide vane has a structural function, in the sense that it allows the passage of forces from the center of the turbomachine towards an outer shroud located in the extension of the fan casing.
    Preferably, the aerodynamic part made in one piece also comprises heat transfer fins arranged in a second internal lubricant cooling passage extending in a second main direction of flow of the lubricant going from the head to the foot. 'blade, said second interior passage being fluidly connected to the first interior passage by an elbow free of heat transfer fins. Thus, the arrangement of the fins in the second interior passage is identical or similar to that in the first interior passage.
    The invention also relates to an aircraft turbomachine, preferably a turbojet engine, comprising a plurality of guide vanes such as those described above, arranged downstream or upstream of a fan of the turbomachine.
    Finally, the subject of the invention is a method of manufacturing such a guide blade, said aerodynamic part of the blade being produced in one piece by additive manufacturing, with the fictitious reference plane of the blade arranged parallel to a support surface of the blade during its manufacture.
    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 view of part of the outlet guide vane shown in the previous figure;
    - Figure 4 corresponds to a sectional view taken along the line IV-IV of Figure 3;
    - Figure 4b is a rear view of the blade shown in Figures 2 to 4a, showing the blade in its position as adopted during its production by additive manufacturing;
    - Figure 5 shows the operation of the heat exchanger constituted by the outlet guide vane shown in the previous figures;
    - Figure 6 is a partial view similar to that of Figure 4b, with the output guide vane being in the form of a second preferred embodiment of the invention; and
    - Figure 7 is a view similar to that of Figure 6, according to an alternative embodiment.
    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. The same is true 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 blower 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 outer 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 stream 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 of the turbojet engine equipment, and / or of the accessories box, and / or of the reducer 20. These vanes 24 thus form part of the fluid circuit (s) in which the lubricant is put into circulation in order to successively lubricate the associated element (s), then to be cooled.
    With reference now to FIGS. 2 to 5, 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 5 can be applied to all the blades 24 of the stator ring centered on the axis 3, or only to certain of these blades.
    The blade 24 may 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 has a foot 34 and a head 36 respectively.
    The foot 34 is used for fixing the blade 24 on the low pressure compressor housing, 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.
    One of the features of the invention resides in the production in one piece of the aerodynamic part 32 of the blade, preferably by additive manufacturing called 3D printing or direct manufacturing. The additive manufacturing of the aerodynamic part 32 is for example carried out by any of the following techniques:
    - selective fusion by laser (from the English “Selective Laser Melting” or “SLM”) or by electron beam (from the English “Electron Beam Melting” or “EBM”);
    - selective sintering by laser (“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.
    The powder used is based on aluminum, or on the basis of another metallic material.
    In addition, the manufacture of the single piece may include the foot 34, and / or the head 36, and / or the platforms 40, without departing from the scope of the invention.
    In this first preferred embodiment of the invention, the aerodynamic part 32 is equipped with two interior passages 50a, 50b substantially parallel to one another, 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, there is provided a second interior passage 50b for cooling the lubricant, 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 outwards 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 be fluidly connected to one the other outside the blade 24, there is for example provided a connecting 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. In this first preferred embodiment as well as in all the other modes, to best optimize the heat exchanges, 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 a leading edge 64. However, an opposite situation can be retained, without departing from the scope of the invention.
    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 zones 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 lower and inner walls 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, as can be seen in FIG. 4a. The maximum thickness of these passages can be of the order of 1 to 2 mm. 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.
    The two interior passages 50a, 50b for cooling the lubricant have the particularity of integrating heat transfer fins. In this first preferred embodiment of the invention, the arrangement and the shape of the fins 80a, 80b are substantially identical or similar in the two passages 50a, 50b. They are also provided in the same densities, although that could be otherwise, without departing from the scope of the invention. Consequently, only the fins 80a, 80b 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 fins of the second interior passage 50b. Furthermore, it is noted that the elbow 54 defines an interior space which is preferably free of fins.
    The fins 80a, 80b in the first passage 50a are substantially parallel to the direction 52a. The fin height in this same direction is of the order of 1 mm, or even less, while their thickness takes a preferably constant value preferably between 0.5 and 1.5 mm. In at least one zone of the passage 50a, and preferably in the entirety of the latter, the fins 80a, 80b are provided in a density for example of about 3 fins / cm 2 . More generally, the density is for example between approximately 1 and 5 fins / cm 2 on average.
    The fins are distributed in rows of fins R1, R2 which follow one another in the first direction 52a, each row being substantially straight and parallel to the transverse direction 60. In the passage 50a, there are thus several tens of rows of fins which follow one another in the first direction 52a. In addition, within each row, the fins 80a, 80b are regularly spaced from one another in the direction 60.
    In this first preferred embodiment of the invention, for two rows Rl, R2 directly consecutive in the first direction 52a, it is provided that within the first row Rl, all the fins 80a are inclined at the same angle acute positive Al with respect to a fictitious plane Pf, represented in FIG. 4b. This plane Pf is here such that the leading edge 64 is inscribed in this plane. In addition, at the inner end of the aerodynamic part 32, the line of rope 33 forms a positive acute angle A4 of a few degrees with the normal 35 in the fictitious plane Pf. The angle A4 is for example between 5 and 15 °. The angle A1 associated with the parallel fins 80a is in the order of 30 to 60 °, and more preferably of the order of 45 °.
    Similarly, it is provided that within the second row R2, all the fins 80b are parallel and inclined by a negative acute angle A2 relative to the fictitious plane Pf, this angle A2 preferably having an absolute value identical to that of angle Al.
    As can be seen in FIGS. 4a and 4b, the views of which are taken in the first direction 52a, the first fins 80a appear at sight arranged between the second fins, and vice versa. The length of the fins being preferably identical or similar within the two rows. These fins then jointly form, in view in the first direction 52a, a broken line whose different successive segments are in planes offset in the direction 52a, these segments being inclined by an angle A3 relative to each other. Here, the angles A3 are of the order of 90 °. However, it is emphasized that this line is only fictitious, since the fins 80a and the fins 80b are in two separate rows offset from each other in the direction 52a. In other words, in a view in direction 52a, the fins 80a, 80b have combined connection points, at the level of which it is also possible to provide connecting spokes (not shown) serving to stiffen the fins in progress of manufacturing, in addition to limiting stress concentrations on the part once in use.
    During the operation shown diagrammatically in FIG. 5, the lubricant 82 circulating through the circuit 56 is introduced into the first interior passage 50a, in the first direction 52a going radially outward. At this stage, the lubricant 82 has a high temperature. A heat exchange then takes place between this lubricant 82 conforming to the fins (not shown in FIG. 5) of the first passage 50a, and the secondary flow 81 conforming to the external surface of the lower and upper surfaces 70, 72 carrying these fins. The lubricant 82, 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 81 and by circulating in the second main direction of flow 52b . Then, the cooled lubricant 82 is extracted from the blade 24, and redirected by the closed circuit 56 to elements to be lubricated and / or to a reservoir of lubricant from which cooled lubricant is pumped to lubricate elements.
    As mentioned above, the aerodynamic part 32 of the blade is made in one piece, by additive manufacturing. The preferred orientation of the blade during its manufacture is shown in Figure 4b. It is such that the fictitious reference plane Pf is arranged parallel to a support surface of the blade provided on a manufacturing frame (not shown). More preferably and as shown in FIG. 4b, this support surface is coincident with the plane Pf, in which the leading edge 64 is inscribed. Consequently, the manufacturing is carried out by superposition of layers in a stacking direction substantially orthogonal to the leading edge. During this manufacture, the fins 80a, 80b can thus be constructed at angles A1 and A2 which guarantee easy manufacture while limiting the length by which they extend between the lower and upper walls 70, 72.
    According to a second embodiment shown in FIG. 6, each row RI, R2 alternately comprises, in the transverse direction 60, fins 80a each forming a positive acute angle A1 with the fictitious reference plane Pf, and fins 80b forming each an acute negative angle A2 with this same plane. Also, instead of incorporating fins all having the same direction of inclination, each row RI, R2 comprises fins arranged alternately with reverse directions of inclination, in order to form broken lines each inscribing in a plane . Preferably, for any two directly consecutive rows RI, R2, the first and second broken lines L1, L2 formed respectively by these rows are offset from one another in the direction 60, and have the same period T. Thus, in a view in direction 52a, the fins 80a of the first row RI cross the fins 80b of the second row R2, and vice versa.
    In the second embodiment shown in FIG. 6, the offset between the two lines L1, L2 in the direction 60, has a value T / 2. The fins of the two consecutive rows are thus in phase opposition, and substantially intersect in their middle.
    According to the alternative represented in FIG. 7, the offset between two directly consecutive rows has a value of T / 3. In view in direction 52a, it is thus possible to see three broken lines L1, L2, L3, formed respectively by three directly consecutive rows RI, R2, RI.
    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-rotary air inlet, the feet of the blades then being connected to this fixed air inlet cone.
    权利要求:
    Claims (11)
    [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 having a first interior passage (50a) for cooling lubricant s extending in a first main direction (52a) of flow of the lubricant going from the foot (34) towards the head (36) the blade, said first interior passage (50a) being partly delimited by a lower surface wall (70 ) and by an upper surface (72) of the blade, characterized in that the aerodynamic part (32) of the blade is made in one piece and also includes heat transfer fins (80a, 80b) arranged in the first passage (50a) by connecting the lower and upper surfaces (70, 72) and extending substantially parallel to the first direction (52a), said fins being distributed in rows of fins following one another in the first main direction (52a) and produced so that for any first (RI) and second row of fins (R2) any directly consecutive, the first row (RI) comprises at least several fins (80a) forming a positive acute angle A1 with a notional dawn reference plane (Pf) parallel to the first direction (52a), while the second row (R2) comprises at least several fins (80b) forming a negative acute angle A2 with said fictitious reference plane (Pf).
    [2" id="c-fr-0002]
    2. Dawn according to claim 1, characterized in that all the fins (80a, 80b) of the blade each form a positive acute angle Al or a negative acute angle A2 with the fictitious reference plane, these angles Al, A2 being between 30 and 60 °.
    [3" id="c-fr-0003]
    3. Dawn according to claim 1 or claim 2, characterized in that for a first (RI) and a second row (R2) of any directly consecutive fins, all the fins (80a) of the first row (RI) form each a positive acute angle A1 with the hypothetical reference plane (Pf), while all the fins (80b) of the second row (R2) each form a negative acute angle A2 with said hypothetical reference plane (Pf).
    [4" id="c-fr-0004]
    4. Dawn according to claim 3, characterized in that the positive acute angle Al is substantially identical for all the fins (80a) of the first row (Rl), while the negative acute angle A2 is substantially identical for all the fins (80b) of the second row (R2).
    [5" id="c-fr-0005]
    5. Dawn according to claim 3 or claim 4, characterized in that the fins (80a) of the first row (Rl) are regularly spaced from each other in a transverse direction (60) of the blade from a leading edge (64) towards a trailing edge (62) of its aerodynamic part (32), in that the fins (80b) of the second row (R2) are regularly spaced from one another in the transverse direction ( 60), and in that in view in the first direction (52a), the fins (80a) of the first row (Rl) are arranged between the fins (80b) of the second row (R2), preferably to form together a broken line.
    [6" id="c-fr-0006]
    6. Dawn according to claim 1 or claim 2, characterized in that for a first (Rl) and a second row of fins (R2) any directly consecutive, each of these rows comprises alternately, in a transverse direction (60 ) from the blade going from a leading edge (64) to a trailing edge (62) of its aerodynamic part (32), fins (80a) each forming an acute positive angle A1 with the fictitious reference plane (Pf) and fins (80b) each forming an acute negative angle A2 with said fictitious reference plane (Pf).
    [7" id="c-fr-0007]
    7. Dawn according to claim 6, characterized in that in view in the first direction (52a), the fins (80a, 80b) of the first row (Rl) jointly form a first broken line (Ll) and the fins ( 80a, 80b) of the second rows (R2) jointly form a second broken line (L2), the first and the second broken line (L1, L2) being offset from each other in the transverse direction (60) of so that at least some fins (80a, 80b) of the first row (RI) cross at least some fins (80a, 80b) of the second row (R2).
    [8" id="c-fr-0008]
    8. Dawn according to claim 6, characterized in that the first (L1) and the second broken line (L2) are periodic with the same period T, and in that they are offset from one another in the direction transverse (60) with a value of T / n, n being a positive integer greater than 1, preferably between two and four.
    [9" id="c-fr-0009]
    9. Directing vane according to any one of the preceding claims, characterized in that the aerodynamic part (32) produced in one piece also comprises heat transfer fins (80a, 80b) arranged in a second interior passage (50b) lubricant cooling extending in a second main direction (52b) of lubricant flow going from the head (36) to the foot (34) the blade, said second inner passage (50b) being fluidly connected to the first passage interior (50a) by an elbow (54) free of heat transfer fins.
    [10" id="c-fr-0010]
    10. 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 .
    [11" id="c-fr-0011]
    11. A method of manufacturing a guide blade (24) according to any one of claims 1 to 9, characterized in that said aerodynamic part (32) of the blade is made in one piece by additive manufacturing, with the dummy reference plane of the blade (Pf) arranged parallel to a support surface of the blade during its manufacture.
    S.62025
    0 OODDDD
    2 ' 6 îA φ ΰ 3
    3/5
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    同族专利:
    公开号 | 公开日
    GB2562360B|2021-11-03|
    US20180258779A1|2018-09-13|
    GB2562360A|2018-11-14|
    US10697312B2|2020-06-30|
    FR3063767B1|2019-04-26|
    GB201803853D0|2018-04-25|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    EP0203458A1|1985-05-15|1986-12-03|Showa Aluminum Corporation|Heat-exchanger of plate fin type|
    US7219720B2|2002-10-11|2007-05-22|Showa Denko K.K.|Flat hollow body for passing fluid therethrough, heat exchanger comprising the hollow body and process for fabricating the heat exchanger|
    US8616834B2|2010-04-30|2013-12-31|General Electric Company|Gas turbine engine airfoil integrated heat exchanger|
    EP2783075A1|2011-11-25|2014-10-01|Siemens Aktiengesellschaft|Airfoil with cooling passages|
    FR2989110B1|2012-04-05|2016-09-09|Snecma|DAWN OF STATOR FORMED BY A SET OF DAWN PARTS|
    US9470095B2|2012-04-24|2016-10-18|United Technologies Corporation|Airfoil having internal lattice network|
    CN103089335A|2013-01-21|2013-05-08|上海交通大学|W-shaped rib channel cooling structure suitable for turbine blade backside cooling cavity|
    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|
    EP3440316A1|2016-05-10|2019-02-13|Siemens Aktiengesellschaft|Ceramic component for combustion turbine engines|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|
    US11149550B2|2019-02-07|2021-10-19|Raytheon Technologies Corporation|Blade neck transition|
    US10871074B2|2019-02-28|2020-12-22|Raytheon Technologies Corporation|Blade/vane cooling passages|
    法律状态:
    2018-02-19| PLFP| Fee payment|Year of fee payment: 2 |
    2018-09-14| PLSC| Publication of the preliminary search report|Effective date: 20180914 |
    2020-02-20| PLFP| Fee payment|Year of fee payment: 4 |
    2021-02-19| PLFP| Fee payment|Year of fee payment: 5 |
    2022-02-21| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1752017A|FR3063767B1|2017-03-13|2017-03-13|OUTPUT DIRECTOR FOR AIRCRAFT TURBOMACHINE WITH IMPROVED LUBRICANT COOLING FUNCTION|
    FR1752017|2017-03-13|FR1752017A| FR3063767B1|2017-03-13|2017-03-13|OUTPUT DIRECTOR FOR AIRCRAFT TURBOMACHINE WITH IMPROVED LUBRICANT COOLING FUNCTION|
    US15/914,150| US10697312B2|2017-03-13|2018-03-07|Outlet guide vane for aircraft turbomachine, with improved lubricant cooling function|
    GB1803853.9A| GB2562360B|2017-03-13|2018-03-09|Outlet guide vane for aircraft turbomachine, with improved lubricant cooling function|
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