• 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 de-icing device for an aeronautical turbomachine separation nozzle (5) (1), comprising a separation nozzle (5) having an outer annular wall (12) delimiting the inside of the flow channel of the secondary flow (4) and an inner annular wall (10) delimiting an inlet of the flow channel of the primary flow (3), and an inner ferrule (14) mounted upstream on the inner annular wall (10) of the spout separation piece and on which are intended to be fixed inlet guide vanes (7), the separating nozzle and the inner ring defining an annular volume (18). The device further comprises an annular baffle (26) positioned within the annular space separating said annular volume into a first annular cavity (28), and a second annular cavity (30) defined between the annular baffle (26) and the outer annular wall (12) of the partition spout. The invention also relates to a fan module comprising such a device and an aviation turbine engine.
    公开号:FR3051016A1
    申请号:FR1654126
    申请日:2016-05-09
    公开日:2017-11-10
    发明作者:Damien Daniel Sylvain Lourit；Francois Marie Paul Marlin
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    Background of the invention
    The present invention relates to the general field of turbomachines. The invention relates more particularly to a de-icing system of an aerospace turbomachine separation spout.
    In an aerospace turbomachine of the double-body and double-flow type, the flow veins of the primary flow and the secondary flow are separated downstream of the fan by a separation nozzle. Within the primary vein, at the inlet of the low pressure compressor (also commonly called "booster"), are a set of fixed inlet guide vanes (also called IGV for "Inlet Guide Vane").
    In certain phases of flight and on the ground, icing atmospheric conditions may be encountered by the turbomachine, especially when the ambient temperature is sufficiently low and in the presence of high humidity. Under these conditions, ice may form on the separation nozzle and the inlet guide vanes. When this phenomenon occurs, it can lead to partial or complete obstruction of the primary vein, and ingestion of loose blocks of ice in the primary vein. An obstruction of the primary vein causes underfeeding of the combustion chamber which may then be extinguished or prevent acceleration of the engine. In the case of detachment of ice blocks, they can damage the compressor downstream and also lead to the extinction of the combustion chamber.
    To prevent the formation of ice on the separator nozzle, techniques are known to collect hot air in the primary stream at a compressor and inject it into the separator nozzle. The hot air injected into the separation spout can then travel through the spout to holes or grooves configured to inject hot air into the primary stream which can also defrost the inlet guide vanes.
    The flow of hot air needed to defrost the separation spout is important. This hot air sampling can reduce the performance and operability of the turbomachine. It would therefore be desirable to be able to increase the efficiency of the deicing of the separation nozzle without increasing the hot air intake in a pressurized part of the turbomachine.
    Object and summary of the invention
    The main purpose of the present invention is therefore to increase the efficiency of the de-icing of the separation spout by proposing a device for de-icing an aeronautical turbomachine separation spout, comprising: a separation spout intended to be positioned downstream a fan of the turbomachine to form a separation between annular flow channels of a primary flow and a secondary flow from the fan, said nozzle having an outer annular wall defining the inside of the flow channel secondary flow and an inner annular wall delimiting an inlet of the flow channel of the primary flow, and an inner ring mounted upstream on the inner annular wall of the separating nozzle and on which are intended to be fixed guide vanes inlet, the separating spout and the inner shell defining an annular volume.
    According to the invention, the deicing device further comprises an annular baffle positioned inside the annular space separating said annular volume into a first annular cavity, and into a second annular cavity defined between the annular baffle and the outer annular wall. separation nozzle.
    In the present description, the terms "upstream" and "downstream" are defined with respect to the direction of flow of the air inside the turbomachine; the terms "inner" and "outer", "axial" and "radial", and their derivatives, are defined with respect to the longitudinal axis of the turbomachine.
    A deicing device according to the invention makes it possible to increase the deicing efficiency of the separation nozzle by reducing in particular the heat losses inside the annular volume defined in the separation nozzle. Indeed, the presence of the annular baffle makes it possible to circulate the hot defrosting air inside the spout in a first cavity having a smaller volume than the total annular volume defined by the separation spout and the inner shell. By reducing this volume, the hot air is channeled and concentrated to the areas of interest to be defrosted, namely the end of the inlet spout and the inner shell. As it is not useful to ensure the defrosting of the outer annular wall of the separator nozzle, the heat losses at this wall are minimized, the hot air no longer circulating directly in the vicinity of the outer annular wall. The deicing efficiency is thus increased without having to modify the defrost hot air flow rate taken from the turbomachine.
    Preferably, the deflector comprises at least one resiliently deformable tab bearing on the inner ring and exerting a force on an upstream end of the baffle in contact with the outer annular wall of the separating spout. The presence of this tab ensures a seal between the baffle and the outer annular wall upstream, and good mechanical strength of the assembly in operation, including ensuring a maintenance of the baffle in position in the annular volume. An elastically deformable tab may for example be metallic, and it can be fixed on the annular baffle by means of rivets. The annular baffle may comprise a plurality of deformable tabs elastically distributed over its entire circumference.
    Also preferably, the baffle is attached downstream on a flange extending from the outer annular wall of the partition spout by means of fixing means. Such fastening means may for example comprise rivets.
    More preferably, the baffle is coated with an insulating layer of heat. Said insulating layer of heat may comprise a material chosen from: a material comprising RTV silicone (for "Room Temperature Vulcanizing"), for example of the DAPCO ™ 2100 type, or a heat-insulating insulation based on compressed silica powder. This insulating heat layer further minimizes heat loss by reducing heat transfer between the first cavity and the second cavity.
    In an exemplary embodiment, the inner ferrule is provided with injection ports configured to inject hot air into the annular flow channel of the primary flow to the inlet guide vanes.
    In an exemplary embodiment, a groove is provided between the inner ferrule and the separating nozzle so that hot air can circulate near one end of the separator nozzle and be injected into the annular flow channel of the flow. primary to defrost the inlet guide vanes.
    In an exemplary embodiment, for easier mounting, the deflector can be divided into a plurality of deflector sectors distributed circumferentially inside the annular volume. For example, the annular baffle may be divided into six angular deflector sectors.
    The de-icing device may further comprise at least one nozzle opening inside the first annular cavity, said nozzle being supplied with hot air by a supply duct connected to a pressurized portion of the turbomachine. The hot air can for example be taken from the high pressure compressor of the turbomachine. For this purpose, the deflector may have at least one boss in which is housed said nozzle. The invention also relates to an aeronautical turbomachine blower module comprising: a blower, a low pressure compressor, inlet guide vanes located upstream of the low pressure compressor and downstream of the blower, and a defrosting device such as that presented above.
    Finally, the invention is an aerospace turbomachine comprising a fan module as defined above.
    BRIEF DESCRIPTION OF THE DRAWINGS Other features and advantages of the present invention will emerge from the description given below, with reference to the accompanying drawings which illustrate an embodiment having no limiting character. In the figures: FIG. 1 is a partial view in longitudinal section of an aeronautical turbomachine equipped with a deicing device according to the invention; FIG. 2 is a first sectional view of a deicing device according to FIG. FIG. 3 is a perspective view of a deflector present in a de-icing system according to the invention, FIG. 4 is a perspective view illustrating the positioning of the deflector in the separation spout, FIG. 5 is a second sectional view of the deicing device, and FIG. 6 is a perspective view of a nozzle that can be used in a deicing device according to the invention.
    Detailed description of the invention
    FIG. 1 partially represents an aeronautical turbomachine 1 of the double-body and double-flow type to which the invention can be applied. In a manner known per se, the turbomachine 1 is axisymmetric with respect to a longitudinal axis XX and comprises an inlet at its upstream end which receives external air, this air supplying a fan 2. Downstream of the fan 2, the The air is distributed between a vein or flow channel of a primary stream 3 (or hot stream) and a vein or flow channel of a secondary stream 4 (or cold stream). These two channels 3, 4 are separated from each other at their inlet by a separation spout 5. Once entered into the flow channel of the primary flow 3, the air then passes through a low pressure compressor 6 ( or "booster"), a high-pressure compressor, a combustion chamber and turbines (the latter elements are not shown in the figures), before being ejected outside the turbomachine. Inner guide vanes 7 or "Inlet Guide Vane (IGV)" are present upstream of the low pressure compressor 6 at the inlet of the flow channel of the primary flow 3.
    As shown in Figures 2 and 5, the partition 5 has a U-shaped or V-shaped longitudinal section rounded at its upstream end. The spout 5 comprises an inner annular wall 10 delimiting the inlet of the flow channel of the primary flow 3, and an outer annular wall 12 delimiting radially inside the flow channel of the secondary flow 4. The outer annular wall 12 has a dimension in the longitudinal direction greater than that of the inner annular wall 10.
    The inner annular wall 10 of the separation spout 5 is extended downstream by an inner shell 14 which carries the inlet guide vanes 7. The inner shell 14 has at an upstream end a hook 16 which allows it to rest on the inner annular wall 10 of the partition 5 being blocked by said wall 10 upstream. A passage, for example a groove, may be provided between the hook 16 and the separation spout 5 to allow hot air to defrost the end of the separation spout 5 and the inlet guide vanes 7. In the illustrated example, the inner ferrule 14 has downstream a radial flange 22 bearing against a flange 24 extending radially from the outer annular wall 12 of the separating nose 5 downstream. In the example illustrated, the inner ferrule 14 may be pierced with injection orifices 20 located downstream of the hook 16 and circumferentially distributed around the inner ferrule 14. The injection orifices 20 may be configured so that operation, hot air is injected into the flow channel of the primary flow 3 to the leading edge of the inlet guide vanes 7 to defrost. The inner and outer annular walls 10 and the inner ferrule 14 together define an annular volume 18 in the separation spout 5.
    According to the invention, an annular baffle 26 is positioned inside the annular volume 18 defined above (FIGS. 2 and 5). The deflector 26 takes the form of a ring which separates the annular volume 18 into a first cavity 28 in which the hot air circulates and, where appropriate, the injection orifices 20 open, and into a second cavity 30 delimited radially. outside by the outer annular wall 12 of the partition spout. The first cavity 28 is thus defined between the inner ferrule 14 and the deflector 26, and more precisely between the inner ferrule 14, the upstream end of the separator nozzle 5, a portion of the outer annular wall 12 and the deflector 26. The Annular volume 18 defined above corresponds to the meeting of the first 28 and second 30 cavities. The deflector 26 comprises, from upstream to downstream: an end 31 in contact with the outer annular wall 12; a portion 32 having a substantially straight longitudinal section; a portion 34 at which the diameter of the deflector 26 increases progressively towards the downstream to accommodate a possible increase in the diameter of the inner ferrule 14; and a flange 36 extending in a radial direction and located at a downstream end of the deflector 26. The flange 36 of the deflector 26 is, in the example illustrated, attached downstream on the flange 24 of the outer annular wall 12 via rivets 38.
    In the example illustrated in FIGS. 2 and 5, the deflector 26 may be covered with a heat-insulating layer 50. This layer 50 also makes it possible to reduce the heat transfer between the first cavity 28 and the second cavity 30 through the As an example, this layer 50 may comprise a material chosen from: a material comprising RTV ("room temperature vulcanizing") silicone, for example of the DAPCO ™ 2100 type, or a heat insulating insulation with base of compressed silica powder.
    It will be noted that the deflector 26 may be divided into a plurality of angular deflector sectors 26, for example six deflector sectors 26, which may be distributed circumferentially in the annular volume 18. An example of an angular deflector sector 26 is shown in FIG. in detail in Figure 3.
    A tab 40 resiliently deformable is present on the deflector 26 and allows to maintain it in position in the annular volume 18. In the example shown, the tab 40 takes the general shape of a spatula. More specifically, the tab 40 comprises a first portion 42 which is fixed on the wall of the deflector 26 opposite the inner ferrule 14, for example by means of rivets 48 (FIG. 3) or by welding, and a second portion 44 provided with a fold 46 at which the tab 40 rests on the inner ring 14. The tab 40 allows to exert on the end 31 of the deflector 26 a radially outward force to maintain the end 31 permanently in contact with the outer annular wall 12. It will be noted that a plurality of tabs 40 may be provided all around the circumference of the deflector 26 to maintain it against the outer annular wall 12. The tabs 40 thus make it possible to minimize even prevent the passage of hot air between the first cavity 28 and the second cavity 30 upstream, while the contact between the flange 36 of the baffle 26 and the flange 24 of the outer annular wall 12 minimizes or even prevents the passage of hot air between the two cavities 28, 30 downstream. For example, a tab 40 elastically deformable can be made of metal material.
    Figure 4 shows the positioning of the deflector 26 inside the annular volume 18. Note that the partition 5 is not shown in this figure. To supply the first cavity 28 with hot air, a supply duct 52 may be provided. This supply duct 52 may for example be connected to a pressurized portion of the turbomachine as a high pressure compressor, on the one hand, and lead into the first cavity 28 via an opening 54 provided in the flanges 22. , 24 and 36, on the other hand. A nozzle 56 connected to the supply duct 52 may be provided in the first cavity 28 to inject hot air into the separation nozzle 5 upstream. A nozzle 56 may be provided with tabs 57 (FIG. 6) pierced to allow its attachment, for example by means of rivets (not shown). Dotted arrows schematize the path of the hot air in Figures 2 and 5.
    In the example illustrated in Figures 3 to 5, to accommodate the presence of the nozzle 56, the deflector 26 may be provided with a boss 58 in which is housed the nozzle 56. Note that a plurality of nozzles 56 can be distributed over the entire circumference of the annular volume 18 of the partition nozzle 5, for example ten nozzles.
    To mount a deicing device according to the invention, comprising in particular the separation spout 5, the inner shell 14 and the annular baffle 26, the annular baffle 26 is first fixed on the flange 24 and then the inner shell 14. When they are used, the nozzles 56 can be mounted in the device before the assembly of the inner ferrule 14 is assembled.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    1. Device for deicing an aeronautical turbomachine separation spout (5) (1), comprising: a separation spout (5) intended to be positioned downstream of a fan (2) of the turbomachine to form a separation between annular flow channels of a primary flow (3) and a secondary flow (4) from the fan, said nozzle having an outer annular wall (12) defining the interior of the flow channel a secondary flow (4) and an inner annular wall (10) delimiting an inlet of the flow channel of the primary flow (3), and an inner ferrule (14) mounted upstream on the inner annular wall (10) of the spout and on which are intended to be fixed inlet guide vanes (7), the separating spout and the inner shell defining an annular volume (18), characterized in that it further comprises an annular baffle ( 26) positioned within the annular volume (18) sep annealing said annular volume into a first annular cavity (28), and a second annular cavity (30) defined between the annular baffle (26) and the outer annular wall (12) of the partition spout.
    [2" id="c-fr-0002]
    2. Device according to claim 1, characterized in that the deflector (26) comprises at least one tab (40) elastically deformable bearing on the inner shell (14) and exerting a force on an upstream end (31) of the deflector in contact with the outer annular wall (12) of the partition spout.
    [3" id="c-fr-0003]
    3. Device according to any one of claims 1 and 2, characterized in that the deflector (26) is attached downstream on a flange (24) extending from the outer annular wall (12) of the spout. separation device (5) by means of fastening means (38).
    [4" id="c-fr-0004]
    4. Device according to any one of claims 1 to 3, characterized in that the baffle (26) is coated with a heat insulating layer (50).
    [5" id="c-fr-0005]
    5. Device according to claim 4, characterized in that said heat-insulating layer (50) comprises a material chosen from: a material comprising RTV silicone, a heat-insulating insulation based on compressed silica powder.
    [6" id="c-fr-0006]
    6. Device according to any one of claims 1 to 5, characterized in that the deflector (26) is divided into a plurality of deflector sectors distributed circumferentially inside the annular volume (18).
    [7" id="c-fr-0007]
    7. Device according to any one of claims 1 to 6, characterized in that it further comprises at least one nozzle (56) opening into the first annular cavity (28), said nozzle being supplied with air hot by a supply duct (52) connected to a pressurized portion of the turbomachine.
    [8" id="c-fr-0008]
    8. Device according to claim 7, characterized in that the deflector (26) has at least one boss (58) in which is housed said nozzle (56).
    [9" id="c-fr-0009]
    Aeronautical turbomachine blower module comprising: a blower (2), a low pressure compressor (6), inlet guide vanes (7) located upstream of the low pressure compressor and downstream of the blower , and a deicing device according to any one of claims 1 to 8.
    [10" id="c-fr-0010]
    10. Aeronautical turbomachine (1) comprising a fan module according to claim 9.
    类似技术:
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    同族专利:
    公开号 | 公开日
    US10494997B2|2019-12-03|
    FR3051016B1|2020-03-13|
    GB2551889A|2018-01-03|
    GB201707252D0|2017-06-21|
    GB2551889B|2021-03-31|
    US20170321604A1|2017-11-09|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
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    US20030035719A1|2001-08-17|2003-02-20|Wadia Aspi Rustom|Booster compressor deicer|
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    法律状态:
    2017-04-13| PLFP| Fee payment|Year of fee payment: 2 |
    2017-11-10| PLSC| Publication of the preliminary search report|Effective date: 20171110 |
    2018-04-23| PLFP| Fee payment|Year of fee payment: 3 |
    2018-09-14| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20180809 |
    2019-04-19| PLFP| Fee payment|Year of fee payment: 4 |
    2020-04-22| PLFP| Fee payment|Year of fee payment: 5 |
    2021-04-21| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1654126A|FR3051016B1|2016-05-09|2016-05-09|DEVICE FOR DEFROSTING A SPOUT FOR AERONAUTICAL TURBOMACHINE|
    FR1654126|2016-05-09|FR1654126A| FR3051016B1|2016-05-09|2016-05-09|DEVICE FOR DEFROSTING A SPOUT FOR AERONAUTICAL TURBOMACHINE|
    GB1707252.1A| GB2551889B|2016-05-09|2017-05-05|A device for de-icing a splitter nose of an aviation turbine engine|
    US15/590,474| US10494997B2|2016-05-09|2017-05-09|Device for de-icing a splitter nose of an aviation turbine engine|
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