• 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 turbojet engine (1) comprising: - a turbine shaft (10), - a fan shaft (20), and - a reduction mechanism (12), coupling the turbine shaft (10). ) and the fan shaft (20), the turbojet engine (1) having a dilution ratio greater than or equal to 10 and the turbine shaft (10) being supported by four bearings (BP # 1, BP # 2, BP # 3, BP # 4) so that the bending deformation modes of the turbine shaft (10) are positioned in transient phase or outside the operating range of the turbojet engine (1).
    公开号:FR3049008A1
    申请号:FR1652176
    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, and a supercritical low pressure shaft, that is to say with a of bending deformation in the operating range.
    BACKGROUND
    A turbofan engine generally comprises, upstream to downstream in the direction of gas flow, a streamlined fan housed in a fan casing, a primary flow annulus and an annular secondary flow space. The air mass sucked by the fan is thus divided into a primary flow, which flows in the primary flow space, and a secondary flow, which is concentric with the primary flow and flows in the flow space. secondary. The primary flow space passes through a primary body comprising one or more stages of compressors, for example a low pressure compressor and a high pressure compressor, a combustion chamber, one or more turbine stages, for example a high pressure turbine, and a low pressure turbine, and a gas exhaust nozzle.
    Typically, the high pressure turbine rotates the high pressure compressor through a first shaft, said high pressure shaft, while the low pressure turbine rotates the low pressure compressor and the blower through a second tree, called low pressure tree. The low pressure shaft is generally housed in the high pressure shaft.
    In order to improve the propulsive efficiency of the turbojet and to reduce its specific consumption as well as the noise emitted by the fan, it has been proposed turbojet engines having a dilution ratio, that is to say the ratio between the flow rate of the flow secondary and primary flow, high. By high dilution rate, here will be understood a dilution ratio greater than 10, for example between 12 and 18.
    To achieve such dilution rates, the blower is decoupled from the low pressure turbine, thereby independently to optimize their respective rotational speed. For example, the decoupling can be performed using a gear such as an epicyclic or planetary reduction mechanism, placed between the upstream end of the low pressure shaft and the blower. The blower is then driven by the low pressure shaft through the reduction mechanism and an additional shaft, said blower shaft, which is fixed between the reduction mechanism and the blower disk.
    This decoupling thus makes it possible to reduce the speed of rotation and the pressure ratio of the fan, and to increase the power extracted by the low pressure turbine.
    The rotational speed of the low-pressure turbine in a turbojet engine comprising a reduction mechanism is therefore much greater than the rotation speed of a low-pressure turbine in a conventional turbojet engine (that is to say without a reduction mechanism ) of equivalent power. The torque to be transmitted by the low pressure shaft to the reduction mechanism is less important than in the case of the conventional turbojet, since the turbojet engines work at equivalent power but the low pressure shaft rotates faster. The low pressure shaft may have a smaller diameter, which facilitates the integration of the high pressure body. However, this reduction in the diameter of the low pressure shaft has the effect of reducing the frequency of eigenmodes, whereas the increase in the speed of rotation of the low pressure shaft enlarges the operating range of the shaft. As a result, the low pressure shaft is caused to exceed a critical speed, which is predetermined, which corresponds to a bending mode of deformation of the shaft in its operating range and to enter into resonance. At resonance, which occurs at the passage of the critical speed of the low pressure shaft, it undergoes overvoltage phenomena that amplify the deformations and forces caused by the unbalance of the tree. Under these conditions, the tree is said to be supercritical.
    A turbomachine rotating at a critical speed in steady state risks a rapid degradation. It is therefore sought to have critical speeds outside the steady state operating ranges of the turbomachine.
    A modal analysis of the architecture of a given low-pressure shaft - which conventionally presents a front and a rear bearing - allows to determine the values of the critical speeds, the shape of the modal deformations as well as the distribution of the deformation energy. between the components of the shaft line: the front, rear bearings and the shaft connecting these bearings. In particular, reference may be made to FIG. 3, which illustrates the modal analysis of a supercritical low-pressure shaft of a high-dilution turbojet according to the prior art. This figure shows, in revolutions per minute, the ground speed (2000 to 4500 rpm), the cruising speed (from 8500 to 9500 rpm), the take-off speed (from 9000 to 9750 rpm). min) for an example of a turbojet having an absolute maximum speed experienced by the low pressure shaft during the entire flight (or "redline" in English) of 10000 rpm. In particular, this figure provides a critical speed value for the first mode of bending deformation of the low-pressure shaft, said mode 1f appears at a critical speed of the order of the order of 5000 rpm. However, in order to limit the area of potential occurrence of instabilities (non-synchronous vibrations), this mode 1f must be higher in frequency. Furthermore, the second mode of bending deformation of the low-pressure shaft, said mode 2f, occurs at a critical speed of the order of 11000 rpm, which is too close to the redline and has the consequence of too much load the motor structure.
    It is therefore necessary to dimension the low pressure shaft in order to push back the deformation modes outside the operating ranges of the turbojet engine, or at least guarantee that they only intervene in the transient phase, and therefore over a period of time. short enough to reduce the risk of damage to the turbojet engine. Indeed, the appearance of such deformation modes can have the effect of preventing the engine revving due to the high deflection of the low pressure shaft and / or generating non-synchronous vibrations, which results in an uncontrolled and divergent increase in the dynamic response of the shaft beyond the critical speed corresponding to the mode 1f.
    For this, it is possible to increase the diameter of the low pressure shaft. However, such an increase is not desirable in a turbojet engine with a high dilution ratio, insofar as it also involves an increase in the bulk of the main body, and therefore a decrease in the dilution ratio.
    SUMMARY OF THE INVENTION
    An object of the invention is therefore to provide a turbofan engine with a high dilution ratio, which has a healthy dynamic situation, that is to say whose modes of deformation appear outside operating ranges, or at least only during transient phases of the turbojet engine.
    For this, the invention provides a turbofan engine comprising: - a turbine shaft, - a fan shaft, and - a reduction mechanism, coupling the turbine shaft and the fan shaft.
    The turbojet engine has a dilution ratio of greater than or equal to 10. In addition, the turbine shaft is supported by four bearings so that the bending deformation modes of the turbine shaft are positioned in transient phase or out of the operating range of the turbojet engine.
    Some preferred but non-limiting characteristics of the turbojet engine described above are the following, taken individually or in combination: the turbojet engine further comprises, from upstream to downstream in the direction of flow of the gases in the turbojet engine: a fan, driven by the fan shaft, a low pressure compressor, driven by the turbine shaft, a high pressure compressor, and a turbine, which drives the turbine shaft in rotation, - the turbojet engine further comprises an inter-compressor housing , extending between the low pressure compressor and the high pressure compressor, and in which a first of the four bearings which supports the turbine shaft is mounted on the inter-compressor casing, the turbojet engine further comprises a sump casing extending between the reduction mechanism and the low pressure compressor, and wherein a second of the four bearings which supports the turbine shaft is mounted on the vein casing, - the first of the four bearings is equipped with a flexible cage and may further comprise an oil film damper, while the second of the four bearings is devoid of a flexible cage, the turbojet comprises in addition, an exhaust casing extending downstream of the turbine, and in which a third of the four bearings is mounted on the exhaust casing; the turbine comprises, from upstream to downstream, a high pressure turbine and a low pressure turbine separated by an inter-turbine casing, the fourth of the four bearings being mounted on the inter-turbine casing, upstream of the third bearing, the fourth of the four bearings is equipped with a flexible cage and may further comprise a an oil film damper, while the third of the four bearings is devoid of a flexible cage, the high pressure compressor is driven by a high pressure shaft, said high pressure shaft being mounted on a front bearing, extending downstream of the first of the four bearings, and a rear bearing, - the high pressure compressor comprises at least eight rotor stages, for example between eight and twelve rotor stages, - an overall compression ratio of the compressor low pressure and the high pressure compressor is greater than or equal to 30, preferably greater than or equal to 40, the dilution ratio of the turbojet engine is between 12 and 18, a reduction ratio of the reduction mechanism is between 2.5 and 5, and / or - an absolute maximum speed experienced by the turbine shaft is between 8,000 rpm and 12,000 rpm, typically around 10,000 rpm.
    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 is a schematic view of an exemplary embodiment of a turbojet according to the invention,
    FIG. 2 is a modal analysis of an exemplary embodiment of a supercritical low-pressure shaft of a turbojet engine with a high dilution ratio according to the invention, and
    FIG. 3 is a modal analysis of a supercritical low-pressure shaft of a turbojet engine with a high dilution ratio according to the prior art.
    DETAILED DESCRIPTION OF AN EMBODIMENT
    In what follows, a turbojet engine 1 will now be described with reference to the appended figures.
    The turbojet engine 1 comprises, in a conventional manner, a fan 2 and a primary body. The primary body comprises, in the direction of gas flow, a low-pressure compressor 3, a high-pressure compressor 4, a combustion chamber 5, a high-pressure turbine 6, a low-pressure turbine 7 and an exhaust nozzle 7 gases.
    The fan 2 comprises a fan disk 2 provided with fan blades 9 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 disk blower 2 is supported by a low pressure shaft 10 which is rotated by the low pressure turbine 7.
    The turbojet engine 1 also comprises an intercompressor casing 11 whose hub is arranged between the low-pressure compressor casing 3 and the casing of the high-pressure compressor 4.
    The turbojet engine 1 has a high dilution ratio, that is to say a dilution ratio greater than or equal to 10, for example between 12 and 18, in order to improve the propulsive efficiency of the turbojet engine 1, to reduce its consumption. specific as well as the noise emitted by the blower 2.
    For this purpose, the fan 2 is decoupled from the low-pressure turbine 7 by means of a reduction mechanism 12. The fan 2 is driven by the low-pressure shaft 10 via an epicyclic reduction gearbox or sun gear, placed between the upstream end of the low pressure shaft 10 and the fan 2, and a fan shaft 20, which is fixed between the reduction mechanism 12 and the fan disk 2.
    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 12 includes an epicyclic reduction mechanism 12.
    The reduction ratio of the reduction mechanism 12 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). .
    The deformation modes of the turbojet engine 1 depend in particular on the dimensioning of the low-pressure shaft 10 and the absolute maximum speed encountered by the low-pressure shaft 10 during the entire flight ("redline", RL).
    The redline RL of the low-pressure shaft 10 is fixed during the production phase of the turbojet engine 1. In this particular case, the redline RL is between 8,000 rpm and 12,000 rpm, typically around 10,000 rpm.
    Conventionally, the length of the low pressure shaft 10 is fixed by the length of the high pressure body, that is to say the length of the high pressure compressor 4, the combustion chamber 5 and the high pressure turbine 6. Here, the high pressure compressor 4 comprises a series of rotating discs (rotor stages), bladed or not, and a series of fixed impellers (rectifier stages). More specifically, the high-pressure compressor 4 comprises at least eight rotor stages, for example between eight and twelve rotor stages.
    Furthermore, the overall pressure ratio ("OPR") of the turbojet engine compressor 1 is at least 30, preferably greater than or equal to 40.
    With respect to the overall pressure OPR, the ratio between the pressure at the inlet of the low-pressure compressor 3 (or "booster" in English) and the pressure at the outlet of the high-pressure compressor 4 will be understood here.
    Thanks to the number of high rotor stages in the high-pressure compressor 4 and the high overall pressure ratio OPR, the turbojet engine 1 compressor has a better fuel efficiency without overloading the booster 3. Such a pressure ratio can in particular be achieved by the reduction mechanism 12 between the fan 2 and the low-pressure turbine 7, which reduces the mass of the turbojet engine 1. The low pressure shaft 10 is centered on the axis of the turbojet engine 1 by a series of bearings. In the case in point, the low pressure shaft 10 is supported by four bearings BP # 1, BP # 2, BP # 3, BP # 4: in this configuration, the deformation modes of the low pressure shaft 10 are displaced transient turbojet 1, with margins of safety compared to stabilized regimes. In particular, reference may be made to FIG. 2, which illustrates the modal analysis of a supercritical low-pressure shaft of a turbojet engine 1 with a high dilution ratio according to the invention, comprising successively four stages BP # 1, BP #. 2, BP # 3 and BP # 4. This figure shows, in revolutions per minute, the ground speed (2000 to 4500 rpm), the cruising speed (from 8500 to 9500 rpm), the take-off speed (from 9000 to 9750 rpm). min) for an example of a turbojet engine 1 having an RL redline of 10000 rpm. Moreover, the first bending deformation mode 1f appears, for this turbojet engine 1 comprising four bearings BP # 1, BP # 2, BP # 3 and BP # 4, at 8000 rpm, while the second mode 2f appears at the redline RL. In one embodiment, the second mode 2f appears beyond 110% of the redline RL, in order to guarantee a margin of safety.
    As can be seen in FIG. 1, the bearing BP # 1, which is situated furthest upstream of the low pressure shaft 10, can be mounted on the one hand on the low pressure shaft 10 and on the other hand on the crankcase. vein 16 which extends between the reduction mechanism 12 and the booster 3.
    The bearing BP # 4, which is located furthest downstream of the low pressure shaft 10, can be mounted on the one hand on the low pressure shaft 10 and on the other hand on the exhaust casing 16 of the turbojet engine 1 .
    The position of the bearings BP # 1 and BP # 4 being conventional, it will not be further detailed in the following.
    The bearing BP # 3, which is adjacent to the bearing BP # 4, can be mounted on the one hand on the low pressure shaft 10 and on the other hand on the interturbine housing 13 (that is to say on the housing extending between the casing housing the high pressure turbine 6 and the casing housing the low pressure turbine 7), upstream of the low pressure turbine 7. In one embodiment, the bearing BP # 3 extends downstream of the bearing HP # 2, which is the most downstream bearing of the high pressure shaft 14.
    The bearing BP # 2, which extends between the bearing BP # 1 and the bearing BP # 3 in the direction of the gas flow in the turbojet engine 1, can be mounted on the one hand on the low pressure shaft 10 and on the other hand on the inter-compressor casing 11, or between the booster 3 and the high-pressure compressor 4. In one embodiment, the BP # 2 bearing extends upstream of the HP # 1 bearing, which is the bearing the most upstream of the high pressure shaft 14.
    As can be seen in FIG. 2, the assembly of the low pressure shaft 10 on four bearings BP # 1, BP # 2, BP # 3 and BP # 4 (rather than two or three bearings, as in the prior art) allows to effectively move the flexural deformation modes 1 f, 2f of the low pressure shaft 10: the mode 1f is positioned in a transient phase of the operating range and with safety margins relative to the stabilized regimes; the 2f mode is positioned outside the operating range and with a comfortable margin compared to the RL redline. In other words, the low pressure shaft 10 remains only a very short time at the critical speed. Typically, the mode 1f can be placed between the ground idle and the cruising / takeoff regimes. At take-off, the turbojet engine shifts from idle speed close to the engine minimum to a take-off regime close to the Redline: the critical speed of the low pressure shaft is therefore likely to appear during the transition between these two regimes.
    It also becomes possible, without risking a flexural deformation mode to appear in steady state, to reduce the diameter of the low pressure shaft 10 and thus the bulk of the primary body to reach, with the reduction mechanism. 12 and the large diameter of the fan 2, a high dilution ratio for the turbojet engine 1. Typically, the low pressure shaft 10 may have an outside diameter of less than fifty millimeters, typically less than forty-five millimeters.
    This positioning of the bearings BP # 1, BP # 2, BP # 3, BP # 4 also makes it possible to reduce the game consumptions (radial displacement) of the booster 3, which is now placed between two levels BP # 2 and BP # 3.
    In one embodiment, the BP # 1 bearing may be devoid of a flexible cage (also known as a squirrel cage) and an oil film compression damper (or "squeeze film damper").
    By oil film compression damper, here will be understood a housing formed in a bearing housing of the corresponding bearing and wherein the outer ring of the bearing is mounted with a small radial clearance. An annular space defined around the ring in this housing is filled with oil and is closed axially by annular sealing elements which are free to rotate in annular grooves of the outer ring of the bearing and which cooperate sealingly with a surface internal cylindrical housing. Oil inlet ports are formed in the housing and open into the aforesaid annular space and oil outlets are formed in the annular sealing members and open out of this annular space to allow to circulate the oil continuously in the annular space and to cool it outside this space in order to evacuate the thermal energy dissipated by the friction resulting from the compression of an oil film by the outer ring of the bearing during its orbital movements in the aforementioned housing. The flexible cage 15 in turn is generally integral with the outer ring of the bearing. In particular, reference may be made to document FR 2 876 758 in the name of the Applicant, which describes an exemplary bearing embodiment comprising an oil film compression damper and a flexible cage. The absence of a flexible cage 15 and oil film compression damper therefore facilitates the integration of the bearing BP # 1, insofar as the space available at the booster 3 is relatively narrow.
    On the other hand, a flexible cage 15 and an oil film compression damper may be placed on the BP # 2 bearing. This flexible cage 15 can be easily integrated on this bearing BP # 2, the space between the inter-compressor casing 11 and the low pressure shaft 10 is greater than that of the booster 3.
    In an alternative embodiment, a flexible cage 15 and an oil film damper may also be placed on the BP # 3 bearing, where the available space is also larger.
    This configuration thus makes it possible to better damp the vibrations of the low-pressure shaft 10, the oil-film damper being more efficient at this position. At the critical speed, the low pressure tree does not orbit at the nodes and its orbitation is maximum at the belly, approaching the compression damper oil flim of the belly, so increases its effectiveness because the clearance in the oil film is more important.
    Optionally, the fourth bearing may also be devoid of a flexible cage 15 and an oil film damper. In this variant embodiment, only the bearings -BP # 2 and BP # 3 are therefore equipped with a flexible cage 15 and an oil film damper.
    权利要求:
    Claims (14)
    [1" id="c-fr-0001]
    A turbofan engine (1) comprising: - a turbine shaft (10), - a fan shaft (20), and - a reduction mechanism (12), coupling the turbine shaft (10) and the fan shaft (20), the turbojet engine (1) having a dilution ratio greater than or equal to 10 and being characterized in that the turbine shaft (10) is supported by four bearings (BP # 1, BP # 2 , BP # 3, BP # 4) so that the bending deformation modes of the turbine shaft (10) are positioned in transient phase or out of the operating range of the turbojet engine (1).
    [2" id="c-fr-0002]
    2. Turbojet engine (1) according to claim 1, further comprising, from upstream to downstream in the direction of flow of the gases in the turbojet engine (1): - a fan (2), driven by the fan shaft ( 20), - a low pressure compressor (3), driven by the turbine shaft (10), - a high pressure compressor (4), and - a turbine (6, 7), which drives the rotation shaft turbine (10).
    [3" id="c-fr-0003]
    3. Turbojet engine (1) according to claim 2, further comprising an inter-compressor casing (11), extending between the low pressure compressor (3) and the high pressure compressor (4), and wherein a first of four Bearings (BP # 2) which supports the turbine shaft (10) is mounted on the inter-compressor housing (11).
    [4" id="c-fr-0004]
    The turbojet engine (1) according to claim 3, further comprising a sump (16) extending between the reduction mechanism (10) and the low pressure compressor (3), and wherein a second of the four bearings ( BP # 1) which supports the turbine shaft (10) is mounted on the vein casing (16).
    [5" id="c-fr-0005]
    5. Turbojet engine (1) according to claim 4, wherein the first of the four bearings (- BP # 2) is equipped with a flexible cage (15) and may further comprise an oil film damper, while the second of the four bearings (BP # 1) is devoid of flexible cage (15).
    [6" id="c-fr-0006]
    6. turbojet engine (1) according to one of claims 4 or 5, further comprising an exhaust casing (16) extending downstream of the turbine (6, 7), and wherein a third of the four bearings ( BP # 4) is mounted on the exhaust casing (16).
    [7" id="c-fr-0007]
    7. A turbojet engine (1) according to claim 6, wherein the turbine (6, 7) comprises, from upstream to downstream, a high pressure turbine (6) and a low pressure turbine (7) separated by an inter-turbine casing. (13), the fourth of four bearings (BP # 3) being mounted on the inter-turbine casing (13), upstream of the third bearing (BP # 4).
    [8" id="c-fr-0008]
    8. Turbojet engine (1) according to claim 7, wherein the fourth of the four bearings (BP # 3) is equipped with a flexible cage (15) and may further comprise an oil film damper, while the third of the four bearings (BP # 4) is devoid of a flexible cage (15).
    [9" id="c-fr-0009]
    9. Turbojet engine (1) according to one of claims 3 to 8, wherein the high pressure compressor (4) is driven by a high pressure shaft (14), said high pressure shaft (14) being mounted on a front bearing ( HP # 1), extending downstream of the first of the four bearings (BP # 2), and a rear bearing (HP # 2).
    [10" id="c-fr-0010]
    10. Turbojet engine (1) according to one of claims 2 to 9, wherein the high pressure compressor (4) comprises at least eight rotor stages, for example between eight and twelve rotor stages.
    [11" id="c-fr-0011]
    11. Turbojet engine (1) according to claim 10, wherein an overall compression ratio of the low pressure compressor (3) and the high pressure compressor (4) is greater than or equal to 30, preferably greater than or equal to 40.
    [12" id="c-fr-0012]
    12. Turbojet engine (1) according to one of claims 1 to 11, wherein the dilution ratio of the turbojet engine (1) is between 12 and 18.
    [13" id="c-fr-0013]
    13. Turbojet engine (1) according to one of claims 1 to 12, wherein a reduction ratio of the reduction mechanism (12) is between 2.5 and 5.
    [14" id="c-fr-0014]
    14. Turbojet engine (1) according to one of claims 1 to 13, wherein an absolute maximum speed (RL) encountered by the turbine shaft (10) is between 8000 revolutions per minute and 12 000 revolutions per minute, typically around 10,000 rpm.
    类似技术:
    公开号 | 公开日 | 专利标题
    FR3049008A1|2017-09-22|TURBOREACTOR COMPRISING A LOWER SUPERCRITICAL PRESSURE TREE
    CA2758175C|2018-04-03|Double-body gas turbine engine provided with an inter-shaft bearing
    EP1626170B1|2007-04-25|Turbomachine with counter-rotating fan
    FR2942273A1|2010-08-20|DOUBLE FLOW MOTOR WITH CONTRAROTATIVE TURBINE WHEELS
    EP3430243B1|2021-01-06|Turbojet engine comprising a simplified bearing lubrication assembly
    FR2866073A1|2005-08-12|TURBOREACTOR HAVING TWO SOLIDARITY CONTRAROTATIVE BLOWERS OF A CONTRAROTATIVE LOW-PRESSURE COMPRESSOR
    FR3048999A1|2017-09-22|TURBOREACTOR LOW GAME BETWEEN BLOWER AND BLOWER HOUSING
    FR3070448B1|2019-09-06|TURBOMACHINE BLOWER RECTIFIER DRAWER, TURBOMACHINE ASSEMBLY COMPRISING SUCH A BLADE AND TURBOMACHINE EQUIPPED WITH SAID DAUTH OR DUDIT TOGETHER
    FR3084909A1|2020-02-14|Efficient gas turbine engine
    WO2017085406A1|2017-05-26|Propulsion unit comprising a main engine and an auxiliary engine
    WO2017158296A1|2017-09-21|Turbojet engine comprising a simplified bearing lubrication unit
    FR3084910A1|2020-02-14|EFFICIENT AIRCRAFT ENGINE
    FR3084908A1|2020-02-14|High propulsion and thermal efficiency aircraft engine
    CA2839248C|2018-12-18|Twin-spool turbine engine design with high-pressure compressor connected to the low-pressure turbine
    FR3068735B1|2019-07-26|TURBOREACTOR WITH LOW NOISE OF BLOW
    EP3580432B1|2021-07-07|Turbojet comprising an architecture of bearings optimised to support a low-pressure shaft
    EP3721093B1|2021-08-25|Pump or compresseur comprising an axial balancing system
    FR3104644A1|2021-06-18|Aeronautical propulsion system with improved propulsive efficiency
    FR3093350A1|2020-09-04|Geared gas turbine engine
    FR3089259A1|2020-06-05|Reduction turbofan
    FR3097595A1|2020-12-25|TURBINE TEMPERATURE SPLITTING IN A GAS TURBINE ENGINE
    WO2021116621A1|2021-06-17|Aero-propulsion system with improved propulsion efficiency
    FR3093349A1|2020-09-04|Geared gas turbine engine
    FR3097597A1|2020-12-25|GAS TURBINE ENGINE WITH HIGH EFFICIENCY BLOWER
    FR3107089A1|2021-08-13|Axial retention of a fan in a turbojet
    同族专利:
    公开号 | 公开日
    FR3049008B1|2018-03-02|
    US20190153978A1|2019-05-23|
    WO2017158294A1|2017-09-21|
    US11047338B2|2021-06-29|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US20040047731A1|2002-09-06|2004-03-11|General Electric Company|Method and apparatus for varying the critical speed of a shaft|
    US20120195753A1|2009-11-20|2012-08-02|Davis Todd A|Gas turbine engine architecture with low pressure compressor hub between high and low rotor thrust bearings|
    EP2360391A1|2010-02-23|2011-08-24|General Electric Company|Epicyclic Gearbox|
    WO2014066815A1|2012-10-26|2014-05-01|United Technologies Corporation|Oil system bearing compartment architecture|
    FR2876758B1|2004-10-19|2008-04-18|Snecma Moteurs Sa|DEVICE FOR SUPPORTING AND GUIDING A ROTATING SHAFT|
    US8402741B1|2012-01-31|2013-03-26|United Technologies Corporation|Gas turbine engine shaft bearing configuration|
    US20130195647A1|2012-01-31|2013-08-01|Marc J. Muldoon|Gas turbine engine bearing arrangement including aft bearing hub geometry|
    US20130192195A1|2012-01-31|2013-08-01|Eric J. Wehmeier|Gas turbine engine with compressor inlet guide vane positioned for starting|
    US10125693B2|2012-04-02|2018-11-13|United Technologies Corporation|Geared turbofan engine with power density range|
    EP3770415A1|2012-10-02|2021-01-27|Raytheon Technologies Corporation|Geared turbofan engine with high compressor exit temperature|
    US10337407B2|2013-03-13|2019-07-02|United Technologies Corporation|Low noise compressor for geared gas turbine engine|
    US20150300264A1|2013-09-30|2015-10-22|United Technologies Corporation|Geared turbofan architecture for regional jet aircraft|
    US20150361878A1|2014-06-13|2015-12-17|United Technologies Corporation|Geared turbofan architecture|GB201704502D0|2017-03-22|2017-05-03|Rolls Royce Plc|Gas turbine engine|
    US11131244B2|2017-11-03|2021-09-28|General Electric Company|Power transmission system and gas turbine engine comprising the same|
    US10808753B1|2019-06-03|2020-10-20|Raytheon Technologies Corporation|Method and apparatus for mounting multiple bearings on a shaft|
    GB201918780D0|2019-12-19|2020-02-05|Rolls Royce Plc|Shaft bearings for gas turbine engine|
    GB201918777D0|2019-12-19|2020-02-05|Rolls Royce Plc|Shaft bearing arrangement|
    GB201918779D0|2019-12-19|2020-02-05|Rolls Royce Plc|Shaft bearings|
    GB201918783D0|2019-12-19|2020-02-05|Rolls Royce Plc|Shaft with three bearings|
    GB201918781D0|2019-12-19|2020-02-05|Rolls Royce Plc|Improved shaft bearing positioning in a gas turbine engine|
    GB201918782D0|2019-12-19|2020-02-05|Rolls Royce Plc|Shaft bearing arrangement|
    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
    FR1652176|2016-03-15|
    FR1652176A|FR3049008B1|2016-03-15|2016-03-15|TURBOREACTOR COMPRISING A LOWER SUPERCRITICAL PRESSURE TREE|FR1652176A| FR3049008B1|2016-03-15|2016-03-15|TURBOREACTOR COMPRISING A LOWER SUPERCRITICAL PRESSURE TREE|
    PCT/FR2017/050596| WO2017158294A1|2016-03-15|2017-03-15|Turbofan comprising a low-supercritical-pressure shaft|
    US16/085,086| US11047338B2|2016-03-15|2017-03-15|Turbofan comprising a low-supercritical-pressure shaft|
    [返回顶部]