• 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 present invention relates to a turbomachine high-pressure turbine assembly (10) comprising at least a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) successively arranged, the rotors ( 12, 14) being mounted on a shaft (2), a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first cavity (C1) disposed between the first rotor (12) and the stator (13), a second cavity (C2) disposed between the stator (13) and the second rotor (14), the assembly being characterized in that it further comprises means (30). , 31) for decreasing the pressure within the first cavity (C1).
    公开号:FR3030614A1
    申请号:FR1462655
    申请日:2014-12-17
    公开日:2016-06-24
    发明作者:Damien Bernard Quelven；Jean-Baptiste Vincent Desforges；Maurice Guy Judet；Ba-Phuc Tang
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] GENERAL TECHNICAL FIELD The present invention relates to a turbomachine high pressure turbine.
    [0002] STATE OF THE ART In order to eject a sufficient quantity of air by mass, an increase in pressure at a constant speed is ensured by the inlet compressor. An important energy release is then caused by the combustion of a fuel, usually kerosene, in the oxygen of the air that passes through the machine. Turbomachines comprise a turbine recovering a portion of the energy released by fuel combustion to drive certain rotating elements including the compressor located at the air inlet. In turbomachines with a double flow of the type shown in FIG. 1, the turbine is generally composed of one or more stages (stator-rotor) at high pressure (HP) and one or more stages at low pressure (BP ).
    [0003] In particular, it can be seen that two-stage HP turbines are among the most efficient. However, a problem specific to this type of HP turbine is the tightness control between the two interdisques cavities. With reference to FIG. 2, which shows a known two-stage HP turbine, we see the first stage formed of a first vane stator 11 (a rectifier) and a first vane rotor 12 (a mobile wheel). and the second stage formed of a second vane stator 13 (another rectifier) and a second vane rotor 14 (another movable wheel). The set of stators 11, 13 forms the fixed blade (the stators 11, 13 are fixed to a casing of the turbomachine 1), and all the rotors 12, 14 form the moving blade (the rotors 12, 14 are mounted on a tree 2). The two inter-disc cavities mentioned are arranged on either side and the second stator 13: upstream of the latter (that is to say between the first rotor 12 and the second stator 13) there is the first cavity C1, and downstream (that is to say between the second stator 13 and the second rotor 14) there is the second cavity C2. The first cavity (or upstream cavity) C1 and the second cavity (or downstream cavity) C2 are separated by a sealing plate (stator / cowling) 20 extending from the platform of the second stator 13 (to which the plate 20 is fixed) to the shaft 2. The plate 20 is a substantially annular plate. In a cooling objective, fluid is injected through the platform of the second stator 13 into the first cavity C1, from which it emerges from a gap between the platforms of the first rotor 12 and the second stator 13. FIG. 2 shows that part of this fluid escapes at the point of junction between the plate 20 and the shaft 2, and rejoins the second cavity C2, from which it emerges from a gap between platforms of the second stator 13 and the second rotor 14. This flow bypasses the second stator 13 and therefore causes a slight loss of efficiency. To combat this phenomenon, it has been sought to improve the seal at the junction point between the plate 20 and the shaft 2 for example via the addition of a labyrinth seal 21. The control of the tightness at this point However, the level is complex, since the plate 20 and the shaft 2 are movable relative to each other. The invention improves the situation by solving in a simple and effective way the problem of sealing between the two interdisques cavities C1 and C2. PRESENTATION OF THE INVENTION The present invention proposes a Turbomachine High Pressure turbine assembly, comprising at least a first bladed rotor, a bladed stator and a second bladed rotor arranged successively, the rotors being mounted on a shaft, a sealing plate extending between the stator and the shaft and separating a first cavity disposed between the first rotor and the stator, a second cavity disposed between the stator and the second rotor, the assembly being characterized in that it comprises besides means for reducing the pressure within the first cavity. According to other advantageous and nonlimiting features: said means for reducing the pressure within the first cavity comprise a centrifugal compressor; said centrifugal compressor comprises a plurality of substantially radial recompression fins extending in the first cavity; said fins are disposed on a downstream face of the first rotor; the fins are disposed at a level of a thinning of the first rotor; the first rotor comprises a plurality of blades, the centrifugal compressor comprising a fin for each blade of the first rotor; the fins have a bent distal end; said means for reducing the pressure within the first cavity further comprise an auxiliary sealing plate disposed in the first cavity facing the centrifugal compressor; the fins and the auxiliary sealing plate have a complementary shape; - The means for reducing the pressure within the first cavity are configured to reduce by at least 50% a pressure difference between the first cavity and the second cavity; the means for reducing the pressure within the first cavity are configured to reduce the pressure difference between the first cavity and the second cavity by at least 90%; said bladed stator is a second stator, the high pressure turbine assembly further comprising a first bladed stator disposed upstream of the first rotor.
    [0004] PRESENTATION OF THE FIGURES Other features and advantages of the present invention will appear on reading the description which follows of a preferred embodiment. This description will be given with reference to the appended drawings in which: FIG. 1 previously described represents a known turbomachine; - Figure 2 previously described is a two-stage high pressure turbine according to the prior art; - Figure 3a shows a detail of a two-stage high pressure turbine according to the prior art; FIG. 3b represents an embodiment of a high pressure turbine according to the invention; - Figures 4a-4b show embodiments of a first rotor of a high pressure turbine according to the invention. DETAILED DESCRIPTION Sealing With reference to FIG. 3a, the permeability of the inter-disk sealing depends largely on the pressure environment prevailing in the interdiscous cavities C1 and C2, which is conditioned by the rate of expansion in the vein at the passage of the second stator 13. In practice, Pavai pressure downstream of this second stator 13 is about 0.4 to 0.8 (advantageously about 0.5) times that upstream of the stator 13. Typically, Pamont and Pavai are of the order of a few bar. This large gap increases the possibility of leaks at the level of the labyrinth seal 21. It is always possible to improve the intrinsic performance of this seal, but aspects such as minimum mounting clearance, integration of NIDDA-type abradable cartridges. or clutter within the inter-disk, that it can never be perfect. The present invention proposes to solve the problems of inter-disk permeability not by further improving the sealing performance at the joint 21, but by acting against the cause of this problem, namely the pressure difference between the two cavities Cl and C2. More specifically, it is proposed a turbomachine high pressure turbine assembly 1, comprising at least a first bladed rotor 12, a bladed stator 13 and a second bladed rotor 14 arranged successively. In practice, the bladed stator 13 is a second stator (as explained in the introduction), the HP turbine 10 further comprising a first stator 11 disposed upstream. By "upstream" or "downstream", we reason according to the direction of the flow of the fluid: leaving the compression chamber, the fluid will pass in order at the first stator 11 and the first rotor 12, then the second stator 13 and finally the second rotor 14. The first stator and rotor 11, 12 constitute a first HP turbine stage 10, and the second stator and rotor 13, 14 constitute a second stage. Preferably the HP turbine 10 is two-stage, but alternatively it may comprise more stages. It will be understood that only the second stator 13 is essential in the context of this assembly of the HP turbine 10, and reference will be made hereinafter in the description as "the stator" 13.
    [0005] The rotors 12, 14 are mounted on a shaft 2. They have a plurality of radial blades extending over their entire circumference, from a platform, which has an inner / outer wall against which the air circulates, defining a vein. In particular, a rotor 12, 14 may be one-piece (and thus support all the blades of the part 1), or formed of a plurality of elementary members each supporting a single blade (a "foot" of the blade) so as to constitute a dawn. We speak in the first case of DAM ("Aubade Monobloc Disk") and in the second case of impeller. In a case as in the dawn, the rotor 12, 14 generally has a thinning 120 under its platform (for mass reduction reasons). In the case of the stators 11 and 13, the blades are fixed to an outer casing. They have a platform defining a radially outer wall of the part 1 (the gas passes inside, the blades extend towards the center) and, if appropriate, a radially inner wall of the part 1 (the gas passes around) in defining a hub, as for a rotor 12, 14. The stators 11, 13 may also be monoblock or not, and fixed or moving blades. Furthermore, a sealing plate 20 extends between the stator 13 and the shaft 2 and separates the first cavity C1 disposed between the first rotor 12 and the stator 13, the second cavity C2 disposed between the stator 13 and the second rotor 14. As explained before, it is a substantially annular plate, generally terminated by a labyrinthine seal 21. And the present HP turbine assembly 10 solves inter-disk permeability problems in that it furthermore comprises means 30, 31 for reducing the pressure within the first cavity C1. Rebalancing the pressures With reference to FIG. 3b, the means 30, 31 for reducing the pressure within the first cavity C1 cause the lowering of the pressure within the cavity C1 of p bar, and to a much lesser extent that of the cavity C2 (of p 'bar, with p >> p'. this description will be assumed l).
    [0006] By repeating the previous examples, respective pressures (when the means 30, 31 are active) of PciP = - upstream - PC2 - Pavai are obtained. X and the pressure differential A between the two chambers C1 and C2 (ie the pressure gradient across the labyrinth seal 21) drops from Ai = Pamont Pavai to - 'C1 PC2 - Pamont Pavai X <Pamont Pavai Preferably the decrease in pressure is such that the pressure delta is substantially reduced, ie the pressures of the first cavity C1 and the second cavity C2 are tended to be equal, ie Δf-> 0, ie X - Pamont Pavai. Thus, the means 30, 31 for reducing the pressure within the first cavity C1 are advantageously configured such that the pressure differential A is at least halved (ie Δf <0.5A ,, that is, ie X> 0.5 x (P - upstream - Pavai)), or at least divided by ten (ie 0.1A ,, that is, X0 x> --- 9 (P upstream upstream Pavai)) . The reduction or even the cancellation of the pressure differential between the first and second cavities C1, C2 reduces or even makes the sealing problem obsolete: if the pressures are equal, there is no flow from the first cavity C1 to the second cavity C2 even though the seal is imperfect. It will be understood that many embodiments of the means 30, 31 for reducing the pressure within the first cavity C1 are possible, starting with mechanical suction within the first cavity C1. Preferably, one uses a "local" centrifugal compressor, as shown in Figure 3b. Centrifugal compressor The centrifugal compressor is a mechanism which drives the fluid in rotation about the axis 2, and consequently causes by centrifugal force its forcible transfer of fluid from the bottom of the cavity C1 upwards (the vein). This thus generates a pressure gradient.
    [0007] More precisely, it is recalled that for a fluid system in rotation, the law of static equilibrium of the pressures is given by the equation: dP - = pWair 2r = pKe2w rotor2r dr With: p: density of the air in kg / m3 r: radius in m wair: speed of rotation of air in rad / s wrotor: speed of rotation of a solid in rad / s Ke: Coefficient of drive = Wair Wrotor The centrifugal compressor 30 creates thus by setting fluid rotation an adverse pressure gradient balancing the centrifugal effect in the first cavity which is maximized when Ke is maximized. Fins Preferably, this compressor 30 comprises a plurality of substantially radial recompression fins 300 extending into the first cavity C1. Insofar as the first rotor 12 already constitutes a solid in rotation with the cavity C1, it It suffices to arrange the fins 300 on a downstream face of the first rotor 12. Thus, the rotation of the rotor 12 naturally and automatically causes the desired effect of reducing the pressure within the first cavity C1. The fins 300 are configured to maximize the Ke coefficient. It should be noted that it is already known to have the recompression fins in a turbine, but never in the first cavity C1, still less on the downstream face of the first rotor 12 and not fortiori for the purpose of reducing the pressure at the first rotor. Within the first cavity C1. For example, the application US Pat. No. 4,759,688 proposes recompression fins facing the upstream face of the first rotor. Such fins (which will be noted that they are integral with the stator) serve only to pump fluid to promote the flow of fluid in the first rotor and thus cooling. They have nothing to do with the present fins 300 which rotate with the first rotor 12. With reference to FIG. 4a which represents the rotor 12, the fins 300 are preferably arranged at the level of the thinning 120 of the first rotor 12. At the end of this thinning, the fins 300 have a curved distal end 301. It is desirable to continue these fins 300 closest to the vein. Such a configuration is preferred in the case of a ventilation said rotor 12 (that is to say, the bottom of the first cavity C1) or ventilation by the stator 13 (as mentioned in the introduction). Indeed the fins 300 are here directly out of the feed holes (for cooling) of the first rotor 12 (next to the labyrinth seal 21). The output radius of the fins 300 is the result of a compromise between mass / strength Mechanical / performance gain.
    [0008] Alternatively, in the case of a so-called "cell bottom" ventilation, that is to say under the platform of the first rotor, it is desirable to position the fins 300 higher than the thinning 120, ie to from the radius of the cell bottom to the vein to keep Ke = 1 during the ascent, in order to limit the losses of entry charges in the "pipes" generated by the fins. In all cases, it is preferable that there are as many fins 300 as blades around the rotor 12 (ie holes if as in Figure 4 the rotor is a wheel adapted to receive the blades, which are separated) . Furthermore it is preferable that the radial passage section is large (limitation of pressure losses), and that the fins 300 are thin (limitation of the addition of mass).
    [0009] It is noted that the addition of fins 300 will reduce the exchange coefficients between the air and the first rotor 12 (Ke is as close as possible to 1, so the relative tangential velocity is almost zero between the disk of the rotor 12 and the fluid), which causes a decrease in the "sensitivity" of the rotor to the air temperature on the downstream face, which makes possible: a control of the thermal of the first rotor 12; a reduction in the axial thermal gradient in the event of any additional recompression fins on its upstream face, hence a drop in disk spillage and improvement of the rotor / stator clearance.
    [0010] Furthermore, the fins 300 increase the first rotor response time 12, which allows a slowing down of the first rotor 12 in the acceleration and deceleration phase, hence a decrease in rotor / stator clearance consumption peaks, and a lowering the associated risk of wear.
    [0011] Auxiliary board To amplify its effect, the compressor 30 (ie the fins 300) can be accompanied by an auxiliary sealing plate 31 disposed in the first cavity C1 facing the centrifugal compressor 30. Such auxiliary plate 31, visible on the FIG. 3b and FIG. 4b decreases the gap, reduces the pressure drops, and makes it easier to make Ke tend towards 1 (since the fluid is forced to rotate at the speed of the first rotor 12).
    权利要求:
    Claims (12)
    [0001]
    REVENDICATIONS1. Turbomachine high pressure turbine assembly (10), comprising at least a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) arranged successively, the rotors (12, 14) being mounted on a shaft (2), a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first cavity (C1) disposed between the first rotor (12) and the stator (13), a second cavity (C2) disposed between the stator (13) and the second rotor (14), the assembly being characterized in that it further comprises means (30, 31) for decreasing pressure within the first cavity (Cl).
    [0002]
    2. The assembly of claim 1, wherein said means (30, 31) for decreasing the pressure within the first cavity (C1) comprises a centrifugal compressor (30).
    [0003]
    The assembly of claim 2, wherein said centrifugal compressor (30) comprises a plurality of substantially radial recompression fins (300) extending into the first cavity (C1).
    [0004]
    4. The assembly of claim 3, wherein said fins (300) are disposed on a downstream face of the first rotor (12).
    [0005]
    5. The assembly of claim 4, wherein the fins (300) are disposed at a thinning (120) of the first rotor (12).
    [0006]
    6. An assembly according to one of claims 4 and 5, wherein the first rotor (12) comprises a plurality of blades, the centrifugal compressor (30) comprising a fin (300) for each blade of the first rotor (12).
    [0007]
    7. Assembly according to one of claims 4 to 6, wherein the fins (300) have a curved distal end (301).
    [0008]
    8. Assembly according to one of claims 2 to 7, wherein said means (30, 31) for reducing the pressure within the first cavity (C1) further comprises an auxiliary sealing plate (31) disposed in the first cavity (C1) facing the centrifugal compressor (30).
    [0009]
    9. An assembly according to claims 3 and 8 in combination, wherein the fins (300) and the auxiliary sealing plate (31) have a complementary shape.
    [0010]
    10. Assembly according to one of claims 1 to 9, wherein the means (30, 31) for reducing the pressure within the first cavity (C1) are configured to reduce by at least 50% a pressure difference between the first cavity (C1) and the second cavity (C2).
    [0011]
    11. The assembly of claim 10, wherein the means (30, 31) for reducing the pressure within the first cavity (C1) are configured to reduce by at least 90% the pressure difference between the first cavity. (Cl) and the second cavity (C2).
    [0012]
    12. An assembly according to one of claims 1 to 10, wherein said bladed stator (13) is a second stator, the high pressure turbine assembly further comprising a first bladed stator (11) disposed upstream of the first rotor ( 12).
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    引用文献:
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    EP0919698A1|1997-11-27|1999-06-02|Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma"|Cooled stator vane|
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    NL192865A|1953-12-02|
    US4759688A|1986-12-16|1988-07-26|Allied-Signal Inc.|Cooling flow side entry for cooled turbine blading|
    JPH0742863U|1993-12-29|1995-08-11|三菱重工業株式会社|Gas turbine structure|
    GB2307520B|1995-11-14|1999-07-07|Rolls Royce Plc|A gas turbine engine|
    CA2263508C|1997-06-19|2003-08-19|Mitsubishi Heavy Industries, Ltd.|Sealing device for gas turbine stator blades|
    US5984636A|1997-12-17|1999-11-16|Pratt & Whitney Canada Inc.|Cooling arrangement for turbine rotor|
    RU2226609C2|2002-06-17|2004-04-10|Открытое акционерное общество "Авиадвигатель"|Turbine of gas-turbine engine|
    FR2892454B1|2005-10-21|2008-01-25|Snecma Sa|DEVICE FOR VENTILATION OF TURBINE DISCS IN A GAS TURBINE ENGINE|
    GB0722511D0|2007-11-19|2007-12-27|Rolls Royce Plc|Turbine arrangement|
    JP2013181577A|2012-02-29|2013-09-12|Mitsubishi Heavy Ind Ltd|Sealing device and rotary machine including the same|
    US9291071B2|2012-12-03|2016-03-22|United Technologies Corporation|Turbine nozzle baffle|
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    法律状态:
    2015-12-11| PLFP| Fee payment|Year of fee payment: 2 |
    2016-06-24| PLSC| Publication of the preliminary search report|Effective date: 20160624 |
    2016-12-02| PLFP| Fee payment|Year of fee payment: 3 |
    2017-11-21| PLFP| Fee payment|Year of fee payment: 4 |
    2018-02-09| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20170717 |
    2018-11-27| PLFP| Fee payment|Year of fee payment: 5 |
    2019-11-20| PLFP| Fee payment|Year of fee payment: 6 |
    2020-11-20| PLFP| Fee payment|Year of fee payment: 7 |
    2021-11-18| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1462655A|FR3030614B1|2014-12-17|2014-12-17|TURBOMACHINE HIGH PRESSURE TURBINE ASSEMBLY|
    FR1462655|2014-12-17|FR1462655A| FR3030614B1|2014-12-17|2014-12-17|TURBOMACHINE HIGH PRESSURE TURBINE ASSEMBLY|
    PCT/FR2015/053597| WO2016097632A1|2014-12-17|2015-12-17|Turbine assembly of an aircraft turbine engine|
    RU2017124883A| RU2705319C2|2014-12-17|2015-12-17|Turbine assembly of aircraft gas turbine engine|
    US15/536,096| US10280776B2|2014-12-17|2015-12-17|Turbine assembly of an aircraft turbine engine|
    JP2017532812A| JP6882979B2|2014-12-17|2015-12-17|Turbine assembly of aircraft turbine engine|
    EP15828746.6A| EP3234309B1|2014-12-17|2015-12-17|Turbine assembly of a turbo-machine of an aircraft|
    CN201580069036.8A| CN107109956B|2014-12-17|2015-12-17|The turbine assembly of aircraft turbine engines|
    BR112017012593-5A| BR112017012593A2|2014-12-17|2015-12-17|aircraft turbomachine turbine set|
    CA2970715A| CA2970715A1|2014-12-17|2015-12-17|Turbine assembly of an aircraft turbine engine|
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