• 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 compressor shell for an aircraft turbomachine, the ferrule (20) being arranged between two bladed rotating wheels (16) and radially to the right of a rectifier (12), the ferrule comprising a sealing device ( 30) comprising one or more wipers (32), including a downstream end wiper (32a) on which there is provided, projecting downstream, a structure (44) for driving and deflecting air designed to axially rectify the leakage air from the end wiper (32a).
    公开号:FR3015591A1
    申请号:FR1363077
    申请日:2013-12-19
    公开日:2015-06-26
    发明作者:Christophe Scholtes；Antoine Robert Alain Brunet；Kevin Eugene Henri Giboudeaux；Hadrien Paul Alexandre Mage
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] The invention relates to the field of aircraft turbine engine compressors. It relates more particularly to the control of the temperature of the sensitive parts of the compressors, and more specifically parts of the high pressure compressor. The invention applies to all types of turbomachines, in particular turbojets and turboprops. STATE OF THE PRIOR ART An aircraft turbomachine high-pressure compressor comprises, alternately in the axial direction, bladed rotating wheels and rectifiers. Conventionally, the rotor of the compressor has a ferrule connecting two successive rotating wheels bladed. This ferrule is located radially to the right of the rectifier arranged between the two wheels, and has sealing wipers cooperating with an abradable coating embedded on the rectifier. The sealing wipers and the abradable coating together form a seal called "labyrinth seal", to significantly reduce the air recirculation within the compressor, this recirculation being of course detrimental to the overall performance of the turbomachine. To reduce this recirculation, it is necessary to create a weak clearance between the stator part and the rotor part. Nevertheless, even if a very low clearance improves the performance and operability of the engine, the integrity of the wipers is often endangered due to their possible contact with the abradable coating, in operation. To reduce this risk, there may be increased play between the stator part and the rotor part, but this lowers the performance and increases the risk of pumping the compressor. Be that as it may, it has been found that there is a hot spot on the ferrule, downstream of the group of wipers, in the direction of flow of the leakage air through these same wipers. The presence of such a hot spot is problematic on the one hand because it can create an accelerated degradation of this part of the ferrule subjected to a very high temperature, and on the other hand because it is likely to cause a thermal gradient generating significant mechanical stresses. In both cases, the mechanical strength of the ferrule is affected.
    [0002] DISCLOSURE OF THE INVENTION The object of the invention is therefore to remedy at least partially the problems mentioned above, encountered in the solutions of the prior art. For this purpose, the subject of the invention is a compressor shell for an aircraft turbomachine, the shell being intended to be arranged between two bladed rotating wheels and radially to the right of a rectifier situated between said two rotating wheels of the compressor, the ferrule comprising a sealing device comprising one or more sealing lips protruding radially outwards towards said rectifier, said sealing device comprising, at its downstream end in a flow direction of the trailing air through this sealing device, an end wiper. According to the invention, said ferrule further comprises, projecting downstream from said end wiper, an air entrainment and deflection structure designed to straighten axially the leakage air resulting from the wiper. end. Also, the invention cleverly to condition the leakage air output of the sealing device, to reduce / eliminate the hot spot observed on the ferrule, downstream of the end wiper. Indeed, studies have shown that in the conventional solutions of the prior art, the observed hot spot is essentially due to a low circumferential speed of the leakage air flow from the end wiper, and to the direction of this leak air flow having a very high radial component. In the invention, the trailing air coming from the end wiper comes into contact with the drive and deflection structure, which makes it possible to benefit from the following advantages. First, the leakage air is deflected to be straightened axially, which prevents it from directly impacting the ferrule in a confined area in which previously was the hot spot. This straightening makes it possible to change the flow of the trailing air, and promotes the contact of the latter with a larger surface of the shell and its surrounding rotating elements. The increase of the contact surface reduces the exchange coefficients and limits the risk of occurrence of a hot spot on the ferrule. In this respect, it is noted that the mechanical stresses are very sensitive to the level and the temperature gradients, especially in the areas for which cracks could be observed in the prior art. With the invention, it has been found that a decrease of one or several tens of degrees, for example from 10 ° C. to 20 ° C., makes it possible to increase the service life of these parts by at least 15%. . In addition, the specific structure of the invention allows to drive the leakage air that comes into contact with it, and therefore increases the coefficient of entrainment. This leads to an increase in the circumferential speed of the air, and therefore also contributes to the reduction of the exchange coefficients.
    [0003] Moreover, it is important to note that the two functions provided by the drive and deflection structure are convergent, since the more the deflection made will lead to an increase in the exchange surface, the higher the speed of the drive. leaking air will tend to increase. In summary, the invention provides for rectifying the flow of leak air and accelerating it, in order to reduce the exchange coefficients and to limit the risks of occurrence of a hot spot on the ferrule. This results in a better mechanical strength of it. Preferably, said drive and air deflection structure takes the form, in longitudinal half-section, of an arm carried by the end wiper.
    [0004] Preferably, the arm is inclined axially so as to extend radially outwardly downstream, in the direction of flow of the leaking air. Preferably, the arm is inclined at an angle of between 20 and 60 ° with respect to the axial direction. Preferably, the air entrainment and deflection structure is formed integrally with the end wiper. Preferably, the end wiper has, from its base, a wiper height H1, and the air entrainment and deflection structure is connected to the end wiper at a base of the wiper. structure whose center is at a height H2 from the base of the end wiper, such that heights H1 and H2 have the following formula: H2> H1 / 3. The subject of the invention is also a compressor for an aircraft turbomachine, comprising two bladed rotating wheels between which a rectifier is placed, the compressor also comprising a ferrule as described above, connecting the two bladed rotating wheels and located radially to the right of the rectifier, the latter comprising an abradable coating cooperating with each wiper of the sealing device fitted to the ferrule. Preferably, the drive and air deflection structure is devoid of contact with the abradable coating. Preferably, the clearance between the abradable coating and the end wiper is less than the minimum distance between the drive and deflection structure, and the abradable coating. Finally, the invention relates to an aircraft turbomachine comprising such a compressor, which is preferably a high pressure compressor. Other advantages and features of the invention will become apparent in the detailed non-limiting description below.
    [0005] BRIEF DESCRIPTION OF THE DRAWINGS This description will be made with reference to the appended drawings among which; - Figure 1 shows a schematic longitudinal sectional view of a turbomachine, according to a preferred embodiment of the invention; - Figure 2 shows a longitudinal half-sectional view of a portion of the high pressure compressor of the turbomachine shown in the previous figure; - Figure 3 shows an enlarged view of a portion of the high pressure compressor of the previous figure; FIG. 4 represents an enlarged view of the compressor, showing in more detail the ferrule specific to the invention; FIGS. 5a and 5b are comparative views schematizing the flow of leakage air with the embodiments of the prior art, and with the design according to the invention shown in the preceding figures; FIG. 6 is an enlarged view of the downstream end wiper equipping the shell, and of the specific air entrainment and deflection structure of the invention; and - Figure 7 is a view similar to that of Figure 4, showing an alternative embodiment. DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS Referring firstly to Figure 1, there is shown an aircraft turbine engine 1, according to a preferred embodiment of the invention. This is a turbojet engine with double flow and double body. Nevertheless, it could be a turbomachine of another type, for example a turboprop, without departing from the scope of the invention. Overall, the turbomachine 1 has a longitudinal axis 3 around which its various components extend. It comprises, from upstream to downstream in a main direction of gas flow through this turbomachine represented by the arrow 5, a fan 2, a low pressure compressor 4, a high pressure compressor 6, a combustion chamber 8, a high-pressure turbine 10 and a low-pressure turbine 12. Referring now to FIGS. 2 and 3, there is shown a portion of the high-pressure compressor 6 in which the invention is located. Nevertheless, it is also applicable to the low pressure compressor 4, without departing from the scope of the invention. The compressor 6 has several stages each comprising a rectifier 12 formed of an annular row of stator vanes carried by a casing 14 of the compressor, and a bladed rotating wheel 16 mounted upstream of its associated rectifier. The two wheels 16 of two directly consecutive stages are connected by a shell 20, for example made in one piece with the disc 22 of one of the two wheels 16. The connection with the other wheel 16 is preferably carried out by bolts 24, which connect the disk 22 of the other wheel 16 with a flange 26 extending the shell 20. The rectifier 12, located between the two wheels 16, is arranged radially right of the shell 20 that surrounds. At its inner annular end, the rectifier 12 comprises an abradable coating 28, preferably of the honeycomb type. This coating cooperates with a sealing device 30 integrated in the ferrule. The device 30 is preferably made in one piece with the ferrule, or may be attached thereto. It is equipped with a plurality of wipers 32 projecting radially outwards towards the abradable coating 28. The wipers 32 are axially spaced from each other, and may be straight as has been shown in FIG. 3, or inclined with respect to the radial direction. Together, the elements 28, 32 form a seal called labyrinth seal, significantly reduce the recirculation of air within the compressor. This recirculation of air corresponds to a leakage air flow 42 passing through the above-mentioned joint, the flow direction of which has been shown schematically by the arrow 40 in FIG. 3. This direction 40 of flow of the leaked air is opposite direction to that of the main direction 5 of gas flow through the turbomachine. In this regard, it is noted that in the rest of the description, the terms "upstream" and "downstream" will be used with reference to the flow direction 40 of the leakage air, and not with reference to the main direction 5 of opposite direction.
    [0006] Referring now to Figure 4, it is shown one of the features of the invention to provide a structure 44 for driving and deflecting leakage air, equipping the wiper 32a located at the downstream end of the sealing device 30. This end wiper 32a is an integral part of the three wipers 32 of the device 30, and corresponds to that located furthest downstream. It is therefore equipped with a structure 44 in the form of arms in the half section of Figure 4, the arm projecting downstream from the end wiper 32a. The arm 44 is designed to be annular and to extend continuously around the axis 5 of the turbojet engine, or to be constituted by elements spaced circumferentially from each other, always around this axis 5. Preferably the arm 44 is made in one piece with the end wiper 32a. The arrows 42 of FIG. 4 show the leakage air flow through the labyrinth seal 28, 32, and then flow downstream after leaving the seal. When it exits the seal by the end wiper 32a, this leakage air flow has a very large radial component, in the direction of the ferrule 20. Nevertheless, this ferrule is not reached directly, since the leakage air flow between first contact with the arm 44. Also, the leakage air is straightened axially by the arm 44. By axial rectification, it is understood that the radial component of the leak air flow is first diminished in favor of its axial component downstream. Preferably, the arm 44 is inclined so that the leakage air flow which escapes there finds a radial component, but this time oriented radially outwards. In addition, the leakage air is rotated by its friction with the outer surface of the rotating arm 44, which increases the drag coefficient. Also, with its increased speed and its reorientation, the leakage air reaches the surrounding rotating elements located downstream of the seal, in particular the rotating wheel 16 and the downstream part of the shell 20. As has been shown schematically in FIG. , the leakage air then flows along substantially concentric flow lines by marrying a large surface of the surrounding elements, namely a very large part of the shell located downstream of the seal 28, 32, and a part of the wheel downstream 16, near the blade roots. This is also visible in FIG. 5b showing the flow 42 of the leakage air, as it could be observed via modeling software. In comparison with FIG. 5a, the flow 42 observed in the prior art is much less extensive, and remains confined in the vicinity of the end wiper 32a, on a restricted axial portion of the ferrule 20 which is the subject of the item. aforementioned hot. The greater amplitude of the flow 42 observed with the present invention makes it possible to increase the coefficient of entrainment, to reduce the exchange coefficients, and thus to limit the risks of occurrence of a hot spot on the ferrule. This results in a better mechanical strength of the ferrule 20. Referring now to FIG. 6, as previously mentioned, it is shown that the arm 44 is inclined with respect to the axial direction so as to extend radially outward downstream. This makes it possible to impose on the leakage air leaving the arm 44 an outwardly directed radial component, as opposed to its inward radial component observed as it exits the end wiper 32a. The angle A of inclination between the arm 44 and the axial direction is for example between 20 and 60.
    [0007] In FIG. 6, other preferred geometrical parameters have been represented. It is first of all the game J1 between the abradable coating 28 and the end of the end wiper 32a. This clearance J1 is less than the clearance J2 between the coating 28 and the end of the arm 44 radially facing the same coating.
    [0008] The ratio between these sets is preferably high, for example greater than five. In the alternative embodiment shown in Figure 7, the arm 44 extends axially downstream beyond the abradable coating 28, for efficiency gain by increasing the air entrainment surface. Nevertheless, here again, the minimum distance J 2 between this arm 44 and the coating 28 remains greater than the clearance 11, preferably in the same proportions as those mentioned above. Therefore, the arm 44 is devoid of contact with the abradable coating 28, even in operation. Returning to FIG. 6, it is shown that the downstream end wiper 32a has, from its base 50 situated in continuity with the outer surface of the ferrule 20, a lip height H1. In addition, the arm 44 is connected to the wiper 32a at a base 52 whose center 54 is at a height H2 from the base 50 of the end wiper 32a. To limit the risk of formation of a recirculation pocket between the arm 44 and the upper part of the end wiper 32a, in which the air leakage could be heated, it is preferably made so that the height H2 is greater than one third of the height Hl. Finally, again for maximum efficiency gain, it is ensured that the axial distance L2 between the distal end of the arm 44 and the center of the end wiper 32a is substantially greater than the axial distance L1 between this end. same center and the axial wall 56 of the coating 28, defining the cavity 58 in which is the end of the wiper 32a. Indeed, the downstream part of the abradable tends to slow the flow in rotation, which is why the end of the arm is located far enough to revitalize the speed of the rotating fluid. Of course, various modifications may be made by those skilled in the art to the invention which has just been described, solely as non-limiting examples.
    权利要求:
    Claims (10)
    [0001]
    REVENDICATIONS1. Ferrule (20) for a compressor (6) for an aircraft turbomachine, the ferrule being intended to be arranged between two bladed rotating wheels (16) and radially to the right of a rectifier (12) situated between said two rotary wheels of the compressor, the ferrule comprising a sealing device (30) comprising one or more sealing lips (32) protruding radially outwardly towards said rectifier (12), said sealing device (30) having, at its end downstream in a direction (40) of leakage air flow through this sealing device, an end wiper (32a), characterized in that said ferrule (20) further comprises, projecting towards the downstream from said end wiper (32a) is an air entrainment and deflection structure (44) for axially straightening the leakage air from the end wiper (32a).
    [0002]
    2. Ferrule according to claim 1, characterized in that said air entrainment and deflection structure takes the form, in longitudinal half-section, of an arm (44) carried by the end wiper (32a). .
    [0003]
    3. Ferrule according to claim 2, characterized in that the arm (44) is inclined axially so as to extend radially outwards in going downstream, in the direction (40) of flow of the flight air.
    [0004]
    4. Ferrule according to claim 3, characterized in that the arm (44) is inclined by an angle (A) between 20 and 60 ° relative to the axial direction.
    [0005]
    5. Ferrule according to any one of the preceding claims, characterized in that the structure (44) for driving and deflecting air is made in one piece with the end wiper (32a) .30
    [0006]
    6. Ferrule according to any one of the preceding claims, characterized in that the end wiper has, from its base (50), a wiping height (H1), and in that the structure (44) of and air deflection is connected to the end wiper (32a) at a base (52) of the structure whose center (54) is at a height (H2) from the base (50) of the end wiper (32a), so that the heights (H1, H2) have the following formula: H2> H1 / 3.
    [0007]
    7. A compressor (6) for an aircraft turbine engine (1), comprising two bladed rotating wheels (16) between which a rectifier (12) is placed, the compressor also comprising a ferrule (20) according to any one of the preceding claims. connecting the two rotating impellers (16) and located radially to the right of the straightener (12), the latter having an abradable coating (28) cooperating with each wiper (32) of the sealing device (30) fitted to the shroud (20). ).
    [0008]
    8. Compressor according to claim 7, characterized in that the structure (44) for driving and deflecting air is devoid of contact with the abradable coating (28).
    [0009]
    9. Compressor according to claim 7 or claim 8, characterized in that the clearance (J1) between the abradable coating (28) and the end wiper (32a) is less than the minimum distance (J2) between the structure ( 44) and the abradable coating (28).
    [0010]
    10. An aircraft turbomachine (1) comprising a compressor (6) according to any one of claims 7 to 9, which is preferably a high pressure compressor. 25 30
    类似技术:
    公开号 | 公开日 | 专利标题
    WO2015092281A1|2015-06-25|Compressor shroud comprising a sealing element provided with a structure for driving and deflecting discharge air
    CA2518355C|2013-03-12|Retention of centring keys of the rings under the variable setting stator blades of a gas turbine engine
    CA2837819C|2019-07-09|Turbomachine element with swirl generating devices
    EP1811131A2|2007-07-25|Set of fixed sectorised diffuser inserts for a turbomachine compressor
    EP2901021B1|2020-05-20|Turbomachine casing and impeller
    WO2015155442A1|2015-10-15|Turbine engine compressor with variable-pitch blades
    WO2010072968A1|2010-07-01|Turbine engine rotor wheel with blades made of a composite material provided with a spring ring
    EP2795068B1|2021-07-14|Turbomachine compressor guide vanes assembly
    EP3069057B1|2019-01-09|Sealing system with two rows of complementary sealing elements
    FR2926612A1|2009-07-24|Rotor drum for e.g. turbo-jet engine of aircraft, has cooling units at internal surface and at right of sealing elements to exchange heat by convection between wall and cooling and ventilation air circulating inside walls and between disks
    WO2009153480A2|2009-12-23|Turbine engine with diffuser
    FR2939852A1|2010-06-18|Stator blade stage for compressor of turboshaft engine e.g. turbopropeller engine, has intermediate blades with axial length or radial height less than that of rectifier blades and extend radially between rectifier blades
    FR3066533B1|2019-06-14|SEALING ASSEMBLY FOR A TURBOMACHINE
    FR2940352A1|2010-06-25|MOBILE WHEEL OF A TURBOMACHINE IN MATERIAL COMPOSITE MATERIAL WITH METAL LECHETTES.
    FR2972482A1|2012-09-14|Sealing sleeve for rotor of e.g. single-stage high pressure turbine of ducted-fan twin-spool turbojet of aircraft, has main body comprising strip cooperating with groove of pin of downstream flange to axially retain flange towards upstream
    EP3420198B1|2020-05-20|Flow-straightener for aircraft turbomachine compressor, comprising air extraction openings having a stretched form in the peripheral direction
    FR2953252A1|2011-06-03|Distribution sector for low pressure turbine of e.g. turbojet of airplane, has outer platform sector comprising stiffeners located in extension of vanes and extended along axis parallel to tangent at upstream and downstream edges of vanes
    FR3026428B1|2019-10-25|RADIANT TURBOMACHINE TURBOMACHINE ROTOR BEARD
    FR3058758A1|2018-05-18|TURBOMACHINE, SUCH AS A TURBO AIRBORNE OR TURBOPROPULSER
    FR3069276B1|2019-08-02|SEALING ASSEMBLY FOR TURBOMACHINE
    FR2933150A1|2010-01-01|Rectifier stage for high pressure compressor of e.g. ducted-fan turbine engine, in aircraft, has stiffening element i.e. sheet metal panel, provided with circular holes and fixed on downstream part and annular edge of outer ferrule
    FR3095472A1|2020-10-30|Turbomachine rotor element
    FR3111677A1|2021-12-24|TURBOMACHINE COMPRESSOR, PROCESS FOR ASSEMBLING THE COMPRESSOR
    WO2016034804A1|2016-03-10|Rotational assembly provided with a wedging device
    FR3085992A1|2020-03-20|TURBINE MOBILE WHEEL BLADE HAVING A CURVILINATED SHAPE
    同族专利:
    公开号 | 公开日
    GB2535126B|2018-08-08|
    US20180163740A1|2018-06-14|
    GB201610630D0|2016-08-03|
    GB2535126A|2016-08-10|
    WO2015092281A1|2015-06-25|
    US10273967B2|2019-04-30|
    FR3015591B1|2016-01-29|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US7540709B1|2005-10-20|2009-06-02|Florida Turbine Technologies, Inc.|Box rim cavity for a gas turbine engine|
    US20090129916A1|2007-11-19|2009-05-21|Rolls-Royce Plc|Turbine apparatus|
    US20110058933A1|2008-02-28|2011-03-10|Mtu Aero Engines Gmbh|Device and method for redirecting a leakage current|
    US20120034072A1|2010-08-03|2012-02-09|Rolls-Royce Plc|Seal assembly|
    US5211533A|1991-10-30|1993-05-18|General Electric Company|Flow diverter for turbomachinery seals|
    US7708520B2|2006-11-29|2010-05-04|United Technologies Corporation|Gas turbine engine with concave pocket with knife edge seal|US10451084B2|2015-11-16|2019-10-22|General Electric Company|Gas turbine engine with vane having a cooling inlet|
    FR3055353B1|2016-08-25|2018-09-21|Safran Aircraft Engines|LABYRINTH SEAL ASSEMBLY FOR TURBOMACHINE COMPRISING ABRADABLE AND INCLINED LECHETTES|
    US10767485B2|2018-01-08|2020-09-08|Raytheon Technologies Corporation|Radial cooling system for gas turbine engine compressors|
    FR3088671B1|2018-11-16|2021-01-29|Safran Aircraft Engines|TIGHTNESS BETWEEN A MOBILE WHEEL AND A TURBOMACHINE DISTRIBUTOR|
    FR3091725A1|2019-01-14|2020-07-17|Safran Aircraft Engines|SET FOR A TURBOMACHINE|
    法律状态:
    2015-12-11| PLFP| Fee payment|Year of fee payment: 3 |
    2016-12-02| PLFP| Fee payment|Year of fee payment: 4 |
    2017-11-21| PLFP| Fee payment|Year of fee payment: 5 |
    2018-02-02| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170719 |
    2019-11-20| PLFP| Fee payment|Year of fee payment: 7 |
    2020-11-20| PLFP| Fee payment|Year of fee payment: 8 |
    2021-11-18| PLFP| Fee payment|Year of fee payment: 9 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1363077A|FR3015591B1|2013-12-19|2013-12-19|COMPRESSOR VIROLE COMPRISING A SEALING LAMINATE EQUIPPED WITH A DRIVING AIR DRIVE AND DEVIATION STRUCTURE|FR1363077A| FR3015591B1|2013-12-19|2013-12-19|COMPRESSOR VIROLE COMPRISING A SEALING LAMINATE EQUIPPED WITH A DRIVING AIR DRIVE AND DEVIATION STRUCTURE|
    GB1610630.4A| GB2535126B|2013-12-19|2014-12-17|Compressor shroud comprising a sealing element provided with a structure for entraining and diverting discharge air|
    US15/104,894| US10273967B2|2013-12-19|2014-12-17|Compressor shroud comprising a sealing element provided with a structure for entraining and diverting discharge air|
    PCT/FR2014/053398| WO2015092281A1|2013-12-19|2014-12-17|Compressor shroud comprising a sealing element provided with a structure for driving and deflecting discharge air|
    [返回顶部]