• 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 turbine rotor for a gas turbine engine, said rotor comprising: - an upstream turbine disk (1); a downstream turbine disk (5); an annular flange (b); a first shell (11) connecting the upstream turbine disk to the annular flange; a second shell (51) connecting the downstream turbine disk to the annular flange; an airflow separator device comprising: a first portion (3), forming a first ring, disposed between the upstream turbine disk and the downstream turbine disk; - A second portion (4), forming a second ring, having a first portion disposed opposite the downstream turbine disk and a second portion disposed between the first ring and the second ring; and a thermal insulation zone (6) disposed between the first portion and the second portion.
    公开号:FR3021348A1
    申请号:FR1454500
    申请日:2014-05-20
    公开日:2015-11-27
    发明作者:Josselin Sicard；Helene Marie Barret；Bertrand Pellaton；Benoit Guillaume Silet
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] TECHNICAL FIELD OF THE INVENTION [0001] The invention relates to a turbine rotor for a gas turbine engine, intended to equip aircraft, and more particularly to a rotor of gas turbine engine. low or medium pressure turbine. STATE OF THE PRIOR ART In turbomachines, it is common to use air taken in particular from the high-pressure compressor to cool the parts located in thermally hot areas, downstream of the combustion chamber of the engine. turbine engine. For example, the rotor of the low pressure turbine must be ventilated with "cool" air to cool the links or fasteners of the vanes on the rotor discs by an appropriate air flow at the connection between the foot vanes and rim of the disc. [0003] FIG. 1 schematically illustrates a prior art turbine rotor comprising an upstream disk 1, a downstream disk 5, an annular flange b. A first ferrule 11 connects the upstream disk 1 to the annular flange b. A second ferrule 51 connects the downstream disk 5 to the annular flange b. The rotor also comprises a flow-separating device 4, a second portion 41 of which is disposed between the first ferrule 11 and the second ferrule 51. These three elements: portion 41, first ferrule 11 and second ferrule 51 are held together by the annular flange. . The flow-separating device is called labyrinth ring, because of its annular shape at 360 ° C and the presence of wipers c. The wipers c of the labyrinth ring 4 make it possible to seal between zones of the turbine under different pressures. They are located vis-à-vis cartridges of abradable material on the stator part. These cartridges prevent the destruction of wipers when they come into contact with the stator. The flow separator device 4 in this rotor has a Y-shape to protect the ferrules of the disks and channel the air flows that cool the disks. Three heat flows f1, f2 and fv coexist within the rotor arrangement: a first flow f1 for the ventilation of the upstream disk, a second flow f2 for the ventilation of the downstream disk and a stream of vein fv coming from a vein d air from the turbine. The first ventilation flow f1, in order to cool the upstream disk, passes (in the direction of the arrow) through the upstream disk by cells formed in the upstream disk 1 and then by at least one hole 45 formed in the flow splitter device 4. [0006] The second ventilation flow f2, in order to cool the downstream disk, passes (in the direction of the arrow) through a plurality of lunules (not visible in FIG. 1) of the device flow separator 4 and through the downstream disc by cavities formed in the downstream disc 5. [0007] The device of FIG. 1 has as a major drawback the presence of thermal gradients at the annular flange due to the cohabitation between the different air flows with different temperatures. The annular flange together holds the ferrule of the upstream disk 11, the ferrule of the downstream disk 51 and the flux separator device 4. The thermal gradients induce mechanical stresses on the annular flange. These mechanical stresses can induce a deterioration or even a rupture of the annular flange. SUMMARY OF THE INVENTION [0008] The invention aims to remedy all or some of the disadvantages of the state of the art identified above, and in particular to provide means for reducing the mechanical stresses at the level of the annular flange. connecting an upstream turbine disk and a downstream turbine disk of a turbine rotor. In this purpose, one aspect of the invention relates to a turbine rotor for a gas turbine engine, said rotor comprising: - an upstream turbine disk; a downstream turbine disc; an annular flange; a first shell connecting the upstream turbine disk to the annular flange; a second ferrule connecting the downstream turbine disc to the annular flange; an airflow separator device comprising: a first portion, forming a first ring, disposed between the upstream turbine disk and the downstream turbine disk; - A second part, forming a second ring, said second portion having a first portion disposed opposite the downstream turbine disk and a second portion disposed between the first ring and the second ring; and a zone of thermal insulation placed between the first part and the second part. Due to this arrangement, the air ventilation flows between the upstream portion and the downstream portion are dissociated. Indeed, the thermal insulation zone as well as the first part and the second part form a physical boundary between the ventilation flow for the cooling of the upstream disk and the ventilation flow for the cooling of the downstream disk. By dissociating the air ventilation flows, the thermal gradient at the flange is reduced or eliminated and thus the mechanical stresses at the flange are reduced or even eliminated. The presence of the thermal insulation zone makes it possible to no longer directly connect, that is to say by the material, the zones in contact with a flow of cold air and a flow of hot air in order to reduce the mechanical stresses due to thermal gradients. In addition to the main features which have just been mentioned in the preceding paragraph, the rotor according to the invention may have one or more additional characteristics among the following, considered individually or according to the technically possible combinations: the insulation zone thermal is a space filled with air; - The thermal insulation zone is disposed between a lower portion of the first portion and an upper portion of the second portion and is opposite the second ferrule; the first part of the flow-separating device and the second part of the flow-separating device are in one piece; the first part of the flow-separating device and the second part of the flow-separating device are separate parts; the first part of the flow-separating device is a labyrinth seal, said labyrinth seal comprising at least one wiper; - A third portion of the first portion bears against the upstream disk, a fourth portion of the first portion bears against the first portion of the second portion, said first portion being configured to radially maintain the first portion. The first part is thus held in abutment between the upstream disk and the second part, the latter part itself being held in abutment against the downstream disk and the annular flange; the annular flange holds between them the first ferrule, the second ferrule and the second portion of the flux separator device. The invention also relates to a turbomachine comprising a rotor according to one of the embodiments described above. The invention also relates to an aircraft comprising a rotor according to one of the previously described embodiments. BRIEF DESCRIPTION OF THE FIGURES [0014] Other features and advantages of the invention will emerge on reading the description which follows, with reference to the appended figures, which illustrate: FIG. 1, a schematic sectional view of a rotor turbine engine for a gas turbine engine according to the prior art; - Figure 2 is a schematic sectional view of a turbine rotor for a gas turbine engine according to one embodiment of the invention. For clarity, identical or similar elements are identified by identical reference signs throughout the figures. DETAILED DESCRIPTION OF AN EMBODIMENT [0016] FIG. 2 is a schematic illustration of a sectional view of a turbine rotor for a gas turbine engine of an aircraft, and more particularly a rotor of a low pressure turbine. The rotor comprises an upstream turbine disk 1 and a downstream turbine disk 5. The upstream turbine disk 1 is part, for example, of the first stage of the low pressure turbine and the downstream turbine disk 5 is part of the second stage of the low pressure turbine. The rotor also comprises a first ferrule 11 and a second ferrule 51. The first ferrule 11 and the second ferrule 51 are cylindrical ferrules. The first ferrule 11 connects the upstream disk 1 to an annular flange b. The second ferrule 51 connects the downstream disk 1 to an annular flange b. The annular flange b makes it possible to maintain in connection the first shell 11 and the second shell 51. [0018] The rotor also comprises an airflow separator device (3, 4).
    [0002] This device has the function of allowing the separation of the air flows circulating in the rotor, namely a first flow f1 (direction of circulation illustrated by an arrow in FIG. 2) which serves for the ventilation of the upstream disk 1 and a second flow f2 (flow direction illustrated by an arrow in Figure 2) which serves for the ventilation of the downstream disk 5. The flux separator device comprises a first portion 3 and a second portion 4. In this embodiment, the first part 3 and the second part 4 are separate pieces. The first part 3 forming a first ring 3 is disposed between the upstream turbine disk 1 and the downstream turbine disk 5. The first part, in this embodiment is a labyrinth seal and comprises at least one wiper c. The wiper c, during operation of the turbine, comes into contact with an abradable material of a cartridge 2 of the stator of the turbine. The second part 4, forming a second ring is disposed between the downstream turbine disk 5 and the first 11 and second ferrule 51. The second portion 4 comprises a first portion 42 disposed opposite the downstream turbine disk 5. The first portion 42 is here in abutment against the downstream turbine disc 5. The second portion 4 comprises a second portion disposed between the first ferrule 11 and the second ferrule 51 and held in position by the annular flange b. The flux separator device also comprises a thermal insulation zone 6 between the first part 3 and the second part 4. In this embodiment, the thermal insulation zone 6 is an air-filled space between the two separate parts. which are the first ring 3 and the second ring 4. The thermal insulation zone 6 is situated between a lower part of the first ring 3 and an upper part of the second ring 4. It is opposite at least the second ferrule 51 which connects the downstream turbine disk 5 to the annular flange b. In this embodiment, it is facing both the second ferrule 51 and the first ferrule 11, the thermal insulation zone 6 is a space filled with air insulating the annular flange 5 of the first ventilation flow f1 and the second ventilation flow f2. As regards the positioning of the first portion 3, a third portion 31 of the first portion bears against the upstream turbine disk 1 and a fourth portion 32 of the first portion bears against the first portion 42 of the second part. The first portion 42 of the second portion radially retains the first portion 3. In this embodiment, the first portion 42 forms a hook in which is inserted the fourth portion 32 of the first portion. The rotor comprises a first ventilation arrangement comprising a plurality of cells (not visible) of the upstream disk 1 and at least one hole 45 of a wall of the first part of the flow-separating device. The first ventilation arrangement allows the circulation of the first ventilation flow f1 for the ventilation of the upstream disk. The first ventilation flow f1 encounters the stream of vein fv coming from an air stream at its exit from the hole 45 made in the wall of the first part of the flow-separating device. The rotor 20 also comprises a second ventilation arrangement comprising a plurality of (non-visible) lunulae formed in the second part of the flow-separating device so as to circulate a second ventilation flow f2 between the first ferrule and the second ferrule towards a space between the second portion 4 of the flow-separating device and the second ferrule 51. The second ventilation arrangement also comprises a plurality of cavities formed in the downstream disk 5. The second ventilation arrangement allows the circulation of the second flow of ventilation f2 for ventilation of the downstream disk. [0022] The invention is not limited to the embodiments previously described with reference to the figures and variants could be envisaged without departing from the scope of the invention.
    权利要求:
    Claims (4)
    [0001]
    REVENDICATIONS1. A turbine rotor for a gas turbine engine, said rotor comprising: an upstream turbine disk (1); a downstream turbine disk (5); an annular flange (b); a first shell (11) connecting the upstream turbine disk to the annular flange; a second shell (51) connecting the downstream turbine disk to the annular flange; an airflow separator device; characterized in that the airflow separator device comprises: - a first portion (3), forming a first ring, disposed between the upstream turbine disk and the downstream turbine disk; - A second portion (4) forming a second ring, said second portion having a first portion disposed facing the downstream turbine disk and a second portion disposed between the first ring and the second ring; and - a thermal insulation zone (6) disposed between the first portion and the second portion.
    [0002]
    2. Rotor according to claim 1 characterized in that the thermal insulation zone is a space filled with air.
    [0003]
    3. Rotor according to any one of claims 1 or 2 characterized in that the thermal insulation zone is disposed between a lower portion of the first portion and an upper portion of the second portion and is opposite the second ferrule.
    [0004]
    4. Rotor according to any one of the preceding claims, characterized in that the first part of the flow-separating device and the second part of the flow-separating device are in one piece. A rotor according to any one of claims 1 to 4 characterized in that the first part of the flow-separating device and the second part of the flow-separating device are separate parts. 6. Rotor according to the preceding claim characterized in that the first part of the flow separator device is a labyrinth seal, said labyrinth seal comprising at least one wiper. 7. Rotor according to any one of claims 4 or 5 characterized in that a first portion of the first portion bears against the upstream disk, a second portion of the first portion bears against the first portion to the second portion, said first portion of the second portion being configured to radially maintain the first portion. 8. Rotor according to any one of the preceding claims characterized in that the annular flange maintains in connection the first ferrule, the second ferrule and the second part of the flow separator device. 9. Turbomachine characterized in that it comprises a rotor according to any one of the preceding claims. 10. Aircraft characterized in that it comprises a rotor according to any one of claims 1 to 8.
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    同族专利:
    公开号 | 公开日
    EP3146157A1|2017-03-29|
    EP3146157B1|2019-07-31|
    US10526893B2|2020-01-07|
    RU2016149668A3|2018-10-24|
    CN106460521A|2017-02-22|
    RU2016149668A|2018-06-20|
    WO2015177429A1|2015-11-26|
    RU2676507C2|2018-12-29|
    CA2949597A1|2015-11-26|
    US20170167264A1|2017-06-15|
    FR3021348B1|2016-06-10|
    CN106460521B|2020-04-07|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    DE1106557B|1957-07-18|1961-05-10|Rolls Royce|Gas turbine, the rotor blades of which have inner cooling ducts|
    US3575528A|1968-10-28|1971-04-20|Gen Motors Corp|Turbine rotor cooling|
    DE3310529A1|1982-03-23|1996-10-31|Snecma|Device for cooling the rotor of a gas turbine|
    EP0169798A1|1984-07-23|1986-01-29|United Technologies Corporation|Rotating seal for gas turbine engine|
    GB2307520A|1995-11-14|1997-05-28|Rolls Royce Plc|Gas turbine engine sealing arrangement|
    EP1264964A1|2001-06-07|2002-12-11|Snecma Moteurs|Arrangement for turbomachine rotor with two blade discs separated by a spacer|
    EP1736635A2|2005-05-31|2006-12-27|Rolls-Royce Deutschland Ltd & Co KG|Air transfer system between compressor and turbine of a gas turbine engine|
    SU1809127A1|1977-07-13|1993-04-15|Motornyj Z|Gas-turbine engine turbine|
    US4526508A|1982-09-29|1985-07-02|United Technologies Corporation|Rotor assembly for a gas turbine engine|
    FR2600377B1|1986-06-18|1988-09-02|Snecma|DEVICE FOR MONITORING THE COOLING AIR FLOWS OF AN ENGINE TURBINE|
    FR2893359A1|2005-11-15|2007-05-18|Snecma Sa|ANNULAR LETTER FOR A LARYRINTH OF SEALING, AND METHOD OF MANUFACTURING SAME|
    FR2937371B1|2008-10-20|2010-12-10|Snecma|VENTILATION OF A HIGH-PRESSURE TURBINE IN A TURBOMACHINE|
    US8382432B2|2010-03-08|2013-02-26|General Electric Company|Cooled turbine rim seal|
    IT1403415B1|2010-12-21|2013-10-17|Avio Spa|GAS TURBINE FOR AERONAUTICAL MOTORS|
    RU2507401C1|2012-11-07|2014-02-20|Российская Федерация, от имени которой выступает Министерство промышленности и торговли Российской Федерации |Gas turbine engine low-pressure turbine|US10415410B2|2016-10-06|2019-09-17|United Technologies Corporation|Axial-radial cooling slots on inner air seal|
    US11098604B2|2016-10-06|2021-08-24|Raytheon Technologies Corporation|Radial-axial cooling slots|
    FR3062414B1|2017-02-02|2021-01-01|Safran Aircraft Engines|MOVABLE RING DRILLING OPTIMIZATION|
    DE102017108581A1|2017-04-21|2018-10-25|Rolls-Royce Deutschland Ltd & Co Kg|Turbomachine with an adaptive sealing device|
    US20210010391A1|2017-12-18|2021-01-14|Safran Aircraft Engines|Damping device|
    FR3075254B1|2017-12-19|2019-11-22|Safran Aircraft Engines|SHOCK ABSORBER DEVICE|
    US10767485B2|2018-01-08|2020-09-08|Raytheon Technologies Corporation|Radial cooling system for gas turbine engine compressors|
    法律状态:
    2015-05-07| PLFP| Fee payment|Year of fee payment: 2 |
    2015-11-27| PLSC| Search report ready|Effective date: 20151127 |
    2016-05-17| PLFP| Fee payment|Year of fee payment: 3 |
    2017-04-13| PLFP| Fee payment|Year of fee payment: 4 |
    2018-02-02| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170719 |
    2018-04-23| PLFP| Fee payment|Year of fee payment: 5 |
    2019-04-19| PLFP| Fee payment|Year of fee payment: 6 |
    2020-04-22| PLFP| Fee payment|Year of fee payment: 7 |
    2021-04-21| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1454500A|FR3021348B1|2014-05-20|2014-05-20|TURBINE ROTOR FOR A GAS TURBINE ENGINE|FR1454500A| FR3021348B1|2014-05-20|2014-05-20|TURBINE ROTOR FOR A GAS TURBINE ENGINE|
    RU2016149668A| RU2676507C2|2014-05-20|2015-05-07|Turbine rotor for gas turbine engine|
    PCT/FR2015/051211| WO2015177429A1|2014-05-20|2015-05-07|Turbine rotor for a gas-turbine engine|
    US15/312,850| US10526893B2|2014-05-20|2015-05-07|Turbine rotor for a gas turbine engine|
    EP15724345.2A| EP3146157B1|2014-05-20|2015-05-07|Turbine rotor for a gas-turbine engine|
    CA2949597A| CA2949597A1|2014-05-20|2015-05-07|Turbine rotor for a gas-turbine engine|
    CN201580029116.0A| CN106460521B|2014-05-20|2015-05-07|Turbine rotor for a gas turbine engine|
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