• 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 turbofan engine (1) having a dilution ratio greater than or equal to 10 and comprising: - a fan (2) comprising a disk (20) provided with vanes (22) at its periphery, a distance (h) between the blade head (22a) (22) and the blower housing (24) being less than or equal to ten millimeters - a concentric primary flow space and a secondary flow space, - a turbine ( 6), housed in the primary flow space and in fluid communication with the blower (2), and - a reduction mechanism (10), coupling the turbine (6) and the blower (2).
    公开号:FR3048999A1
    申请号:FR1652162
    申请日:2016-03-15
    公开日:2017-09-22
    发明作者:Jeremy Dievart；Yanis Benslama；Nathalie Nowakowski
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    FIELD OF THE INVENTION The invention relates to the field of turbomachines, and more particularly to turbofan engines having a high or very high dilution ratio.
    BACKGROUND
    As visible in Figs. 4 and 5 appended, a turbofan engine 1 generally comprises, from upstream to downstream in the direction of the flow of gas, a fan 2 housed in a housing 24 of a fan. The fan 2 comprises a disk 20 of a fan (or rotor) provided with vanes 22 at its periphery which, when they are rotated, cause an air flow in the turbojet engine 1. The mass of air sucked by the blower 2 is divided into a primary flow, which flows in a primary flow space, and a secondary flow, which is concentric with the primary flow and flows in a secondary flow space. The primary flow space passes through a primary body comprising one or more stages of compressors, for example a low pressure compressor 4 and a high pressure compressor 3, a combustion chamber, one or more turbine stages, for example a high turbine. pressure 5 and a low pressure turbine 6, and a gas exhaust nozzle 7.
    Typically, the high-pressure turbine 5 rotates the high-pressure compressor 3 via a first shaft, said high-pressure shaft 5a, while the low-pressure turbine 6 rotates the low-pressure compressor 4 and the fan 2 via a second shaft, said low pressure shaft 6a. The low pressure shaft 6a is generally housed in the high pressure shaft 5a, said shafts 5a, 6a being fixed to the structural parts (including the inlet casing, which comprises a fixed blade wheel which supports the fan casing) of the turbojet engine 1 via bearings, typically downstream of the separation nozzle 8 configured to separate the primary flow and the secondary flow.
    When the rupture of a blade 22 of fan 2 ("fan blade out" or FBO, in English terminology), for example under the impact of a foreign body, the fan 2 undergoes significant unbalance. In particular, reference may be made to FIG. 1, which illustrates the response of the blower 2 to an unbalance resulting from the loss of a blade 22 as a function of the rotational speed (in revolutions per minute) of the low pressure shaft 6a. . The bending deformation mode M 'of the fan shaft 2 being situated in the operating range of the turbojet engine 1, the turbojet engine is highly susceptible to damage.
    It has therefore been proposed to place a decoupler 9 between the blower 2 and the inlet casing hub, typically at the separation spout 8, in order to allow the blower 2 to operate in degraded mode despite the presence of an unbalance important. For this purpose, when the breaking load of the decoupler 9 is reached due to a loss of fan blade, one of the bearing links 26 is broken (generally the connection of the bearing 26 before the shaft 2a supporting the blower 2), which makes it possible to change the dynamic situation of the fan 2 and to reduce the mode of deformation in bending of the fan shaft 2 towards the low frequencies. After rupture of this connection, the rotor 20 of the fan 2 can then refocus around its new axis of inertia, which induces a reduction of the forces transmitted to the structures. The distance H between the blades 22 of the fan 2 and the casing of the fan 2 must then be sufficiently large (of the order of about forty millimeters) between the head 22a of the blades 22 of the fan 2 and the housing 24 of blower to allow free orbiting of the rotor 20 and prevent the blades 22 of the fan 2 come into contact with the fan casing 24 (see Figure 2a). It is therefore possible to realize fan casings 2 of reduced mass, since the latter must not hold the contact forces with the blades in case of FBO. However, such an increase in the diameter of the fan casing 24 implies an increase in the size of the nacelle, which has a negative impact on the turbojet 1's drag, its mass and its specific consumption.
    There are also motors without decoupler 9 in which the fan rotates freely during a blade loss and is driven only by the flow of the air flow (phenomenon of autorotation, or "windmilling" in English). ). In this type of engine, in the case of FBO, the head 22a of the vanes of the fan 2 abuts against the part facing the casing of the fan 2 (due to the unbalance). The contact zone is, however, designed to withstand the contact forces and friction (see Figure 2b), as the engine transmits more effort to the structure of the aircraft. In this way, the fan can rotate freely in degraded mode until the landing of the aircraft. In order to guarantee that the forces of contact with the fan blades 22 by the fan casing 24 in degraded mode are maintained, the external surface of the fan casing 24 is structurally reinforced by means of reinforcements 28 which generally comprise an annular flange. reported and fixed on its external face. Such a reinforcement 28, however, substantially increases the mass of the fan casing 24 and therefore of the turbojet engine 1, thus increasing the specific consumption of the turbojet engine 1.
    Consequently, none of the proposed solutions makes it possible to obtain a turbojet engine 1 of reduced mass whose fan can freely rotate in degraded mode in the event of rupture of a fan blade 22 (FBO).
    SUMMARY OF THE INVENTION
    An object of the invention is to provide a turbofan engine which has a reduced mass in comparison with conventional turbojet turbofan engines while being able to run in degraded mode in case of rupture of a fan blade.
    For this, the invention proposes a turbofan engine comprising: a fan housed in a fan casing, said fan comprising a disk provided with fan blades at its periphery, each blade of a fan comprising a blade head extending at a distance from the disk, - a concentric primary flow space and a secondary flow space, - a turbine, housed in the primary flow space and in fluid communication with the fan, and - a reduction mechanism , coupling the turbine and the fan.
    The turbojet has a dilution ratio greater than or equal to 10. Furthermore, a distance between the head of the fan blades and the fan casing is less than or equal to ten millimeters.
    Certain preferred but non-limiting characteristics of the turbojet engine described above are the following, taken individually or in combination: it has a dilution ratio of between 12 and 18, the distance between the head of the fan blades and the fan casing; is less than or equal to six millimeters, preferably equal to between five and six millimeters, - a thickness of the fan casing is less than or equal to fifteen millimeters, preferably less than or equal to twelve millimeters, for example less than or equal to ten millimeters - An outer diameter of the blower is between eighty inches (203.2 centimeters) and one hundred inches (254.0 centimeters), preferably between eighty inches (203.2 centimeters) and ninety inches (228.6 centimeters). Optionally, the diameter of the blower is between eighty inches (203.2 centimeters) and ninety inches (228.6 centimeters) and a thickness of the blower housing is between nine millimeters and twelve millimeters, for example equal to ten millimeters, - a difference in the thickness of the fan casing, between an upstream end and a downstream end of the said fan casing, is less than or equal to ten millimeters, - the fan casing is made of a composite material comprising a densified fibrous reinforcement by a matrix, said fibrous reinforcement comprising fibers chosen from the following group: carbon, glass, aramid, silica carbide and / or ceramic, and / or said matrix comprising a polymer chosen from the following group: epoxide, bismaleimide and / or polyimide, the fan casing has a thickness of between eight and twenty millimeters, preferably between ten and eighteen meters, for example between twelve millimeters and fifteen millimeters, the reduction mechanism is epicyclic or planetary and has a reduction ratio of between 2.5 and 5, and / or the turbojet engine further comprises a separation nozzle extending downstream. of the blower and configured to separate the primary flow space and the secondary flow space, said turbojet engine being devoid of decoupler between the blower and said separation spout.
    BRIEF DESCRIPTION OF THE DRAWINGS Other features, objects and advantages of the present invention will appear better on reading the detailed description which follows, and with reference to the appended drawings given by way of non-limiting examples and in which:
    FIG. 1 illustrates the loading (in Newton N) applied by the disk of the blower of a turbojet according to the prior art to the front bearing of the blower 2 before the decoupler breaks according to the speed of the low shaft. pressure 6a (in revolutions per minute (rpm)),
    FIG. 2a is a schematic partial view of the front of a first example of a turbojet according to the prior art, on which are notably visible a fan, a decoupler and structural parts of the turbojet,
    FIG. 2b is a schematic partial view of the front of a second example of a turbojet according to the prior art, the turbojet being devoid of a decoupler,
    FIG. 3 illustrates the loading (in Newton N) applied by the disk of the fan of a turbojet according to an embodiment of the invention on the front bearing of the fan as a function of the speed of the low pressure shaft ( in revolutions per minute (rpm),
    FIG. 4 is a schematic partial view of the front of an exemplary embodiment of a turbojet according to the invention, on which are notably visible a fan and structural parts of the turbojet, and
    Figure 5 is a partial sectional view of an exemplary embodiment of a turbojet according to the invention.
    DETAILED DESCRIPTION OF AN EMBODIMENT
    In what follows, a turbojet engine 1 will now be described with reference to FIGS. 3 to 5 attached.
    The turbojet engine 1 comprises, in a conventional manner, a fan 2 housed in a fan casing 24, an annular primary flow space and an annular secondary flow space. The primary flow space passes through a primary body. The primary body having been described above, it will not be further detailed here.
    The fan 2 comprises a fan disc 20 provided with fan blades 22 at its periphery which, when they are rotated, cause the flow of air into the primary and secondary flow spaces of the turbojet engine 1. The rotor 20 is driven by the low pressure shaft 6a, which is centered on the X axis of the turbojet 1 by a series of bearings and is rotated by the low pressure turbine 6.
    The fan casing 24 is of generally annular shape and has an inner face extending facing the blades 22 of the fan 2, and an outer face, opposite to the inner face and extending opposite the nacelle. An abradable material 25 may be attached to the inner face of the fan casing 24, facing the fan blades 22.
    In order to improve the propulsive efficiency of the turbojet engine 1, to reduce its specific fuel consumption as well as the noise emitted by the fan 2, the turbojet engine 1 has a bypass ratio (in English terminology), which corresponds to the ratio between the airspeed. secondary flow (cold) and the flow rate of the primary flow (hot, which passes through the primary body). High dilution rate, here will be understood a dilution ratio greater than 10, for example between 12 and 18. At this In effect, the blower 2 is decoupled from the low-pressure turbine 6, thus making it possible to independently optimize their respective rotational speed, For example, the decoupling can be carried out using a gearbox, such as a reduction mechanism. 10 epicyclic ("star gear reduction mechanics" in English language) or planetary gear ("planetary gear reduction mechanics" in Anglo-Saxon), placed between the upstream end (by supply to the flow direction of the gases in the turbojet engine 1) of the low pressure shaft 6a and the fan 2. The fan 2 is then driven by the low pressure shaft 6a via the reduction mechanism 10 and an additional shaft, said fan shaft 2a, which is fixed between the reduction mechanism 10 and the disk 20 of the fan 2.
    This decoupling thus makes it possible to reduce the rotational speed and the pressure ratio of the fan 2 ("fan pressure ratio" in English terminology) and to increase the power extracted by the low pressure turbine 6.
    To calculate the dilution ratio, the flow rate of the secondary flow and the flow rate of the primary flow are measured when the turbojet engine 1 is stationary in a standard atmosphere (as defined by the manual of the International Civil Aviation Organization (ICAO) Doc 7488/3, 3rd edition) and at sea level.
    In one embodiment, the reduction mechanism 10 comprises an epicyclic reduction mechanism.
    The reduction ratio of the reduction mechanism 10 is preferably between 2.5 and 5.
    The diameter of the blower 2 may be between eighty inches (203.2 centimeters) and one hundred inches (254.0 centimeters), preferably between eighty inches (203.2 centimeters) and ninety inches (228.6 centimeters). . By diameter of the blower 2, here will be understood the radial distance between the axis X of revolution of the turbojet engine 1 and the head 22a of the blades 22 of the fan.
    The Applicants have realized that, thanks to the reduction mechanism 10 which makes it possible to reduce the speed of rotation of the fan 2 (of the order of 30% relative to the speed of rotation of the fan of an equivalent turbojet engine without reduction mechanism) and to stiffen the disk 20 of the fan 2 (the shaft 2a being short and on two supports 26), the distance h between the head 22a of the fan blades 22 and the fan casing 24 is significantly reduced when a blower blade rupture 22 (FBO). Therefore, it is possible to reduce the distance h between the head 22a of the fan blades 22 and the fan casing 24 so that said distance h is at most equal to ten millimeters, or even less than six millimeters, for example of order of five to six millimeters. It will be noted that a distance h of five to six millimeters between the head 22a of the fan blades 22 and the inner face of the fan casing 24 corresponds generally to the conventional thickness of the abradable material 25 which is fixed on the fan casing 24. .
    Due to this short distance h, the external diameter of the fan casing 24 is therefore smaller than in the case where a large distance is left in order to prevent any contact between the fan blades 22 and the fan casing 24, which also makes it possible to limit the dimensions of the nacelle in which the fan casing 24 is housed, and therefore the turbojet 1 drag.
    A distance h of the order of five to six millimeters makes it possible to guarantee, for a turbojet engine 1 with a high dilution ratio, that the fan blades 22 do not come into contact with the fan casing 24 in the event of dawn rupture. 22 (FBO). It is therefore no longer necessary for the fan casing 24 to play a role of retention of the blades 22 in the event of rupture of the fan blade 22, which makes it possible to dispense with the reinforcements usually fixed on the external face of the casing. 24 fan (especially the reinforcements 28, which are visible in Figure 2b of the prior art). Thus, the fan casing 24 may, for example, have a thickness e which is less than or equal to fifteen millimeters in the zone extending facing the fan blades 22, preferably less than or equal to thirteen millimeters, typically less than ten millimeters in the case of a fan casing 24 made of a metallic material. By thickness e, here will be understood the dimension extending between a lower face of the fan casing 24 (on which can be fixed an abradable material 25) and an outer face (which extends opposite the nacelle).
    For example, for a fan diameter 2 between eighty inches (203.2 centimeters) and ninety inches (228.6 centimeters), the thickness e of the fan casing 24 is preferably between nine millimeters and twelve millimeters. typically equal to ten millimeters.
    For a fan diameter 2 between ninety inches (228.6 centimeters) and one hundred inches (254.0 centimeters), the thickness e of the fan casing 24 is preferably between twelve millimeters and fifteen millimeters.
    In addition, the difference in thickness e between the outer face and the inner face of the fan casing 24, along the fan casing 24 (that is to say along the X axis of the turbojet engine 1), can be at most ten millimeters. This is especially allowed by the absence of reinforcements 28 on the outer face of the fan casing 24.
    Since the fan casing 24 is no longer capable of holding the support forces of the fan blades 22 in the case of FBO, it also becomes possible to produce the fan casing 24 in a composite material of the matrix-reinforced fibrous reinforcement type. Such a material thus makes it possible to greatly reduce the mass of the fan casing 24, and therefore the mass and the specific consumption of the turbojet engine 1. The fiber reinforcement comprises fibers configured to form the reinforcement of the composite material and take up most of the efforts mechanical. The fibers may in particular be based on carbon, glass, aramid, silicon carbide and / or ceramic. As for the matrix, its main purpose is to transmit the mechanical forces to the reinforcement, to ensure the protection of the reinforcement with respect to the various environmental conditions and to give the desired shape to the product produced. For example, the matrix may comprise a polymer, in particular of the epoxide, bismaleimide or polyimide type. In this case, the blower housing 24 may for example have a thickness e of less than or equal to ten millimeters, when the blower 2 has an outer diameter of about 210 centimeters.
    The Applicants have furthermore established that, in a turbojet engine 1 with a high dilution ratio, the decoupler situated between the bearings 28 and the separation nozzle 8 can be eliminated without impairing the recovery of the forces resulting from a break in the blade. blower (FBO). Indeed, thanks to the reduction mechanism 10 which reduces the speed of rotation of the fan 2 in operation and the length of the shaft 2a which directly drives the fan 2 in rotation, the modes of deformation of the bearings 26 which the shaft 2a of the fan are pushed out of the operating ranges of the turbojet engine 1. In particular, reference may be made to FIG. 3, which illustrates the flexural deformation mode M of the fan 2 with respect to the absolute maximum speed. RL encountered by the fan shaft 2a during the entire flight ("redline" in English). In particular, since the mode of deformation M to FBO is beyond the redline RL, it is outside the operating range of the turbojet engine 1. The loading of the fan blade 20 transmitted by the fan shaft 2a to the bearings 26 of the fan 2 is therefore greatly reduced and thus makes it possible to break one of the bearing links 26 of the fan 2.
    In the example illustrated in Figure 3, the redline RL of the fan shaft 2a is between 2000 revolutions per minute and 4000 revolutions per minute, typically around 3000 revolutions per minute.
    The suppression of the decoupler thus contributes to reducing the mass of the turbojet engine 1, and thus makes it possible to improve the specific fuel consumption of the turbojet engine 1.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    1. turbofan engine (1) comprising: - a fan (2) housed in a fan casing (24), said fan (2) comprising a disc (20) provided with blades (22) of a fan at its periphery. , each fan blade (22) having a blade head (22a) extending away from the disk (20), - a concentric primary flow space and a secondary flow space, - a turbine (6) , housed in the primary flow space and in fluid communication with the blower (2), and - a reduction mechanism (10), coupling the turbine (6) and the blower (2), the turbojet engine (1) having a dilution ratio greater than or equal to 10 and being characterized in that a distance (h) between the head (22a) of the fan blades (22) and the fan casing (24) is less than or equal to ten millimeters.
    [2" id="c-fr-0002]
    2. Turbojet engine (1) according to claim 1, having a dilution ratio between 12 and 18.
    [3" id="c-fr-0003]
    3. turbojet engine (1) according to one of claims 1 or 2, wherein the distance (h) between the head (22a) of the blades (22) of the fan and the housing (24) blower is less than or equal to six millimeters, preferably equal to between five and six millimeters.
    [4" id="c-fr-0004]
    4. Turbojet engine (1) according to one of claims 1 to 3, wherein a thickness (e) of the fan housing (24) is less than or equal to fifteen millimeters, preferably less than or equal to twelve millimeters, for example lower than or equal to ten millimeters.
    [5" id="c-fr-0005]
    5. Turbojet engine (1) according to one of claims 1 to 4, wherein an outer diameter of the fan (2) is between eighty inches (203.2 centimeters) and one hundred inches (254.0 centimeters), preferably between four twenty inches (203.2 centimeters) and ninety inches (228.6 centimeters).
    [6" id="c-fr-0006]
    6. Turbojet engine (1) according to one of claims 1 to 5, wherein a thickness difference (e) of the fan housing (24) between an upstream end and a downstream end of said fan housing (24), is less than or equal to ten millimeters.
    [7" id="c-fr-0007]
    7. Turbojet engine (1) according to one of claims 1 to 6, wherein the fan housing (24) is made of a composite material comprising a fiber reinforcement densified by a matrix, said fibrous reinforcement comprising selected fibers in the group next: carbon, glass, aramid, silica carbide and / or ceramic, and / or said matrix comprising a polymer selected from the following group: epoxide, bismaleimide and / or polyimide.
    [8" id="c-fr-0008]
    8. Turbojet engine (1) according to claim 7, wherein the fan housing (24) has a thickness of between eight and twenty millimeters, preferably between ten and eighteen millimeters, for example between twelve millimeters and fifteen millimeters.
    [9" id="c-fr-0009]
    9. Turbojet engine (1) according to one of claims 1 to 8, wherein the reduction mechanism (10) is epicyclic or planetary and has a reduction ratio of between 2.5 and 5.
    [10" id="c-fr-0010]
    10. Turbojet engine (1) according to one of claims 1 to 9, further comprising a separation nozzle (8) extending downstream of the fan (2) and configured to separate the primary flow space and the secondary flow space, said turbojet engine (1) being devoid of decoupler between the fan (2) and said separation nozzle (8).
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    同族专利:
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    EP3430241A1|2019-01-23|
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    CN108779683B|2022-02-11|
    CN108779683A|2018-11-09|
    WO2017158295A1|2017-09-21|
    US20190085790A1|2019-03-21|
    引用文献:
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    EP2623723A2|2012-02-06|2013-08-07|United Technologies Corporation|Clearance control system for a gas turbine engine section|
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    GB201820934D0|2018-12-21|2019-02-06|Rolls Royce Plc|Low fan noise geared gas turbine engine|
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    GB201820945D0|2018-12-21|2019-02-06|Rolls Royce Plc|Low noise gas turbine engine|
    US20200347736A1|2019-05-03|2020-11-05|United Technologies Corporation|Gas turbine engine with fan case having integrated stator vanes|
    CN112412878A|2020-09-30|2021-02-26|航天材料及工艺研究所|Capturing type ballistic impact resistant composite material fan casing and manufacturing method thereof|
    法律状态:
    2017-03-08| PLFP| Fee payment|Year of fee payment: 2 |
    2017-09-22| PLSC| Publication of the preliminary search report|Effective date: 20170922 |
    2018-02-20| 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 |
    2020-02-20| PLFP| Fee payment|Year of fee payment: 5 |
    2021-02-19| PLFP| Fee payment|Year of fee payment: 6 |
    2022-02-21| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1652162A|FR3048999B1|2016-03-15|2016-03-15|TURBOREACTOR LOW GAME BETWEEN BLOWER AND BLOWER HOUSING|
    FR1652162|2016-03-15|FR1652162A| FR3048999B1|2016-03-15|2016-03-15|TURBOREACTOR LOW GAME BETWEEN BLOWER AND BLOWER HOUSING|
    EP17714868.1A| EP3430241A1|2016-03-15|2017-03-15|Turbofan|
    US16/085,441| US20190085790A1|2016-03-15|2017-03-15|Gas turbine engine with minimal tolerance between the fan and the fan casing|
    CN201780018019.0A| CN108779683B|2016-03-15|2017-03-15|Gas turbine engine with minimum tolerance between fan and fan housing|
    PCT/FR2017/050597| WO2017158295A1|2016-03-15|2017-03-15|Turbofan|
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