• 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 turbomachine comprising: • a fan rotor supplying a low-pressure compressor primary stream and a secondary stream with a hub of diameter D1; • a reducer interposed between the fan rotor and a turbine shaft; a low pressure compressor housed in a housing of external diameter D2, characterized in that - the diameter of the fan rotor is greater than 82 inches (2.08 meters), - the blower pressure ratio is between 1.10 and 1.35 - The pitch diameter of the ring gear of the gearbox is between 0.15 and 0.35 times the diameter of the rotor of the fan, and the outer diameter D2 is greater than the diameter D1 of the hub.
    公开号:FR3065994A1
    申请号:FR1753856
    申请日:2017-05-02
    公开日:2018-11-09
    发明作者:Kevin Morgane LEMARCHAND；Gilles Alain Marie Charier；Nathalie Nowakowski；Nicolas Jerome Jean Tantot
    申请人:Safran Aircraft Engines SAS；
    IPC主号:
    专利说明:

    TURBOMACHINE WITH BLOWER ROTOR AND REDUCER DRIVING A LOW PRESSURE COMPRESSOR SHAFT
    GENERAL TECHNICAL AREA AND PRIOR ART
    The present invention relates to the field of turbomachinery.
    The search for the minimization of polluting emissions linked to air transport involves in particular the improvement of all the efficiencies of the propulsion systems, and more particularly of the propulsive efficiency which characterizes the efficiency with which the energy which is communicated to the air passing through the engine is converted into useful thrust force.
    The elements influencing this propulsive efficiency in the first order are those linked to the low pressure parts of the propulsive system, which immediately contribute to the generation of the thrust: low pressure turbine, low pressure transmission system, blower rotor and secondary channel guiding the flow of the latter.
    The known guiding principle for improving the propulsive efficiency consists in reducing the compression ratio of the fan, thereby reducing the flow speed at the outlet of the engine and the losses by kinetic energy which are linked to it. One of the main consequences of this reduction in flow speed at the outlet from the engine is that it is necessary to have the low pressure part (secondary flow) treat a greater mass air flow in order to ensure a level given thrust, fixed by the characteristics of the aircraft: this therefore leads to an increase in the dilution rate of the engine.
    This increase in secondary flow has the direct effect of requiring an increase in the diameter of the fan, and therefore in the external dimensions of the surrounding retaining casing, as well as of the nacelle constituting the aerodynamic envelope of the casing in question. The question of the ability to integrate propulsion systems of increasingly large dimensions under an aircraft wing then arises with increasing acuity, in a context where ground clearance is limited.
    In addition to the dimensional aspects, the increase in the dilution rate strongly penalizes the mass of the propulsive system, in particular via a very significant increase in the mass of the fan casing, dimensioned for centrifugal retention in the event of blading ejection.
    It is thus noted that the highest dilution rates, although synonymous with the best propellant yields, are accompanied by penalties of mass, drag and difficulties of installation under wing so great that most of the gain thus hoped for finds itself overshadowed by these highly penalizing elements.
    An alternative to this paradigm consists in getting rid of the concept of fairing of the low pressure part: the propulsive architecture thus constituted bears the name of turboprop (case of a single low pressure rotor not faired, qualified as propeller), or "open rotor" according to the English terminology commonly used (case of two low pressure contra-rotating rotors, qualified as counter-rotating propellers). This alternative architecture, if it makes it possible to overcome the constraints of mass and drag of friction of the fairing of the abutment now nonexistent, however poses other problems: first, the absence of fairing makes the context of certification very different vis-à-vis the potential ejection of the blade from the low pressure rotor, and requires consideration of more complex technologies on the rotor (blade called "fail safe" according to English terminology, for example); secondly, the absence of fairing around the low pressure part makes the aerodynamic operation of its rotor very sensitive to variations in flight conditions (in particular speed), and limits the maximum flight speed admissible by the aircraft. Finally, the absence of fairing induces a much lower specific flow rate than a faired solution, leading, for a given level of thrust, to much larger external dimensions than a faired solution, thus aggravating the difficulty of installation on the airplane cell in dimensional terms.
    The complete removal of the fairing of the low pressure part, if it appears to be an acceptable option for aircraft of moderate dimensions and flying at low to medium speeds (regional / short haul applications), seems to induce too many disadvantages for use. on higher thrust classes (medium - long haul), for which the flight speed capacity is a little negotiable expectation of the operators.
    OVERVIEW OF THE INVENTION
    A general aim of the invention is to propose a solution making it possible to overcome the drawbacks of the techniques of the prior art.
    In particular, an object of the invention is to propose a solution making it possible to ensure the efficiency of the low pressure turbine coupled to a slow fan rotor, with a low pressure ratio.
    According to one aspect, the invention proposes a turbomachine comprising
    • a blower rotor supplying an air flow to a primary stream of low pressure compressor and a secondary stream, comprising a hub of diameter DI • a reducer interposed between the blower rotor and a turbine shaft of the low pressure compressor housed in a casing with external diameter D2, characterized in that - the diameter of the blower rotor is greater than 82 inches (2.08 meters), - the pressure ratio of the blower is between 1.10 and 1.35, - the pitch diameter of the crown of the reducer is between 0.15 and 0.35 times the diameter of the blower rotor, and the external diameter D2 being greater than the diameter DI of the hub.
    Such a reducer ensures the efficiency of the low pressure turbine.
    Its sizing and positioning avoid clutter which would be prohibitive.
    It allows integration compatible with a hub ratio allowing the necessary mach / flow rate blowers.
    In another aspect, the turbine includes a compact nacelle.
    More particularly, the nacelle is a protective fairing surrounding the rotor of the blower, said fairing having a length limited to said rotor.
    In this way, the nacelle is reduced to its simplest expression, the fairing only conferring on it an aerodynamic boundary function around the fan rotor.
    It extends neither upstream nor downstream of said rotor and consists of an aerodynamic fairing dimensioned to provide protection against the ejection of the blades.
    The function of guiding the flow upstream of the blower rotor (role previously played by the air intake) is eliminated.
    It is also the same for the functions of control of the downstream pressure field (role previously played by the secondary nozzle) and realization of the reverse thrust function (role previously played by a specific system integrated into the nacelle).
    Furthermore, the blades are advantageously of variable pitch.
    This allows the blower operating point to be controlled according to the flight conditions. Note that the association of a blower with a very low pressure ratio and the absence of a secondary nozzle (one of whose key roles is to control the position of the fan's operating point in its field) induces high variability in fan operating lines between low altitude and high altitude conditions; potentially, this situation induces a difficulty in operability (lack of pumping margin) under ground conditions, and a device for controlling the operating point of the fan according to the flight conditions proves necessary.
    PRESENTATION OF THE FIGURES Other characteristics and advantages of the invention will emerge further from the description which follows, which is purely illustrative and not limiting, and must be read with reference to the appended FIG. 1, which is a schematic representation in sectional view (half-view) illustrating the integration of a fan reducer in a turbomachine according to a possible embodiment of the invention.
    DESCRIPTION OF ONE OR MORE MODES OF IMPLEMENTATION AND IMPLEMENTATION
    The turbomachine T illustrated in FIG. 1 has a streamlined blower architecture with a very high dilution rate (so-called UHBR architecture or “Ultra High By-Pass Ratio” according to the Anglo-Saxon terminology generally used).
    It includes a nacelle 1, a fan rotor 2, as well as a primary stream 3, defined in a casing 5.
    FIG. 1 also shows an intercompressor casing 8 of the turbomachine, an inter-turbo casing 9, as well as an exhaust casing 10.
    The nacelle 1 is compact and in particular of reduced length. In particular, it does not incorporate an air inlet or a secondary nozzle upstream or downstream of the blower.
    It also does not incorporate a reverse thrust mechanism.
    Its main functions are to ensure the aerodynamic fairing of the turbomachine and the retention of the blades / blades of the fan and is only designed for this purpose.
    A rectifier 4 is interposed between the nacelle 1 and the casing 5 and makes it possible to hold said nacelle 1
    In one possible embodiment, part of the nacelle 1 can be made common with an already existing surface on the aircraft, such as for example the lower surface of the wing.
    The rotor blades 2 of the fan are blades 6 with variable setting (mechanism 6a).
    The setting of the blades 6 can in particular be controlled to control the fan in operation. The very low pressure ratio of the latter in fact induces variations in the cycle parameters between the ground and flight conditions of an unusual magnitude, in particular with regard to the HP turbine operating temperatures and the nozzle expansion rates. .
    The blade timing control allows adaptation to these differences in operating conditions.
    It is also used to brake the aircraft or to contribute to it.
    The diameter D3 of the rotor 2 of the fan is large: greater than 82 inches (2.08 meters), and preferably between 90 (2.29 meters) and 150 inches (3.81 meters).
    The blower pressure ratio (FPR or Fan Pressure Ratio according to English terminology) is low: between 1.10 and 1.35.
    Given this dimensioning, the speed of rotation of the rotor 2 is low.
    Consequently, a reduction gear 7 is provided for driving the shaft 8 of the low pressure turbine.
    This reducer 7 allows a high low pressure turbine speed: between 3.5 and 8 times the speed of the rotor 2 and preferably between 5 and 6 times the speed of the latter. The reduction ratio and the torque to be transmitted define the size of the reduction gear. Here the reducer 7 is of the epicyclic type and therefore its reduction ratio is defined by: 1+ (the number of teeth of the crown / the number of teeth of the central sun gear). The torque to be transmitted defines the minimum size of the teeth and the minimum diameter of the central planetary or here the power of the reducer must be between 10 and 40 MW. The pitch diameter of the crown D4 is therefore complex to integrate for such a reduction ratio and is between 0.15 and 0.35 times the diameter of the fan.
    The diameter D3 of the fan is determined in a conventional manner, by projection of the radial component at the head of a fan blade 6, on a radial straight line passing through the leading edge of the blade, at the foot of that -this.
    The hub ratio is defined as the ratio between the internal radius at the foot of the fan blade 6, measured at the leading edge of the blade (at its design setting, in the case where the blade has a variable setting ), and the outer radius of the leading edge of the blade 6 projected on the same line. To guarantee a good performance of the turbomachine, the hub ratio is limited as much as possible, so the diameter of the hub is between 0.25 and 0.35 the diameter of the fan.
    In particular, the radius at the bottom of the fan can be between 300 and 600 mm.
    To integrate a reducer with a high reduction ratio while keeping the hub ratio as small as possible, the inventors have found that it was possible to make a protrusion on the casing surrounding the reducer without harming the aerodynamic characteristics of the primary vein. However, this protuberance must be limited, there is thus a relationship between the external diameter (diameter D2 of the gearbox housing) and the diameter of the fan hub which is greater than 1 (D2> D1) and between 1 and 1.10, and preferably less than 1.04.
    Such a ratio allows both the aerodynamic shape desired for the primary stream 3 and the integration of the servitudes of the reducer (evacuation of the oil, for example) and of the blower (pitch change system), while maintaining a lowest possible hub ratio. The inlet housing in which the reducer is integrated is particularly congested because it must support the reducer, take up the axial thrust generated by the blower via the ball bearing and support the low pressure shaft.
    Furthermore, the input power of the reducer is between 10 and 40 MW (at takeoff (@ T / O or "Take -Off" according to English terminology - altitude 0, Mach between 0.15 and 0.28)).
    The propulsion system thus formed meets the following objectives: - maximization of the propulsive efficiency thanks to the blower with very low pressure ratio; - competitiveness in terms of fuel consumption for thrust classes and flight speed of medium-haul and larger applications (thrust> 15,000 Ibf in take-off condition 0 m / zero speed / ISA conditions; 0.65 <Mach of flight cruise <0.9).
    The fairing allows a minimum drag and is not penalizing in mass.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    1. Turbomachine comprising • a fan rotor supplying air flow to a primary stream of low pressure compressor and a secondary stream, comprising a hub of diameter DI • a reducer interposed between the fan rotor and a turbine shaft of the low compressor pressure housed in a casing with external diameter D2, characterized in that - the diameter of the blower rotor is greater than 82 inches (2.08 meters), - the pressure ratio of the blower is between 1.10 and 1.35, - the pitch diameter of the reducer crown is between 0.15 and 0.35 times the diameter of the blower rotor, and the external diameter D2 being greater than the diameter DI of the hub.
    [2" id="c-fr-0002]
    2. Turbomachine according to claim 1, characterized in that the diameter of the fan rotor is between 90 inches (2.29 meters) and 150 inches (3.81 meters).
    [3" id="c-fr-0003]
    3. Turbomachine according to one of the preceding claims, characterized in that it comprises a nacelle which is a protective fairing surrounding the rotor of the fan, said fairing having a length limited to said rotor.
    [4" id="c-fr-0004]
    4. Turbomachine according to one of the preceding claims, characterized in that the blades of the fan rotor are of the variable pitch type.
    [5" id="c-fr-0005]
    5. Turbomachine according to one of the preceding claims, characterized in that the reduction gear is epicyclic.
    [6" id="c-fr-0006]
    6. Turbomachine according to one of the preceding claims, characterized in that the fan hub ratio is between 0.25 and 0.35.
    [7" id="c-fr-0007]
    7. Turbomachine according to one of the preceding claims, characterized in that the reduction ratio of the reducer is between 3.5 and 8.
    [8" id="c-fr-0008]
    8. Turbomachine according to one of the preceding claims, characterized in that the reduction ratio of the reducer is of the order of 5 or 6.
    [9" id="c-fr-0009]
    9. Turbomachine according to one of the preceding claims, characterized in that the ratio between the external radius of the reduction gear and the radius at the bottom of the fan is between 1 and 1.10.
    [10" id="c-fr-0010]
    10. Aircraft comprising a turbomachine according to one of the preceding claims.
    类似技术:
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    同族专利:
    公开号 | 公开日
    US20200080496A1|2020-03-12|
    EP3619417A1|2020-03-11|
    US11268450B2|2022-03-08|
    WO2018202962A1|2018-11-08|
    FR3065994B1|2019-04-19|
    CN110651112A|2020-01-03|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    WO2014143248A1|2013-03-15|2014-09-18|KARAM, Michael|Ultra high bypass ratio turbofan engine|
    WO2015012923A2|2013-05-09|2015-01-29|United Technologies Corporation|Turbofan engine front section|
    US5281094A|1991-05-13|1994-01-25|Alliedsignal Inc|Electromechanical apparatus for varying blade of variable-pitch fan blades|
    US6231301B1|1998-12-10|2001-05-15|United Technologies Corporation|Casing treatment for a fluid compressor|
    EP2971596B1|2013-03-10|2020-07-15|Rolls-Royce Corporation|Gas turbine engine and corresponding method|
    GB2515747A|2013-07-01|2015-01-07|Techtronic Floor Care Tech Ltd|Surface cleaning apparatus|
    JP6554282B2|2014-12-24|2019-07-31|川崎重工業株式会社|Aircraft engine equipment|
    FR3034130B1|2015-03-25|2018-04-06|Safran Aircraft Engines|SOUFFLANTE BLADE DISASSEMBLY|
    GB201703521D0|2017-03-06|2017-04-19|Rolls Royce Plc|Geared turbofan|
    FR3080886B1|2018-05-02|2020-10-30|Safran Aircraft Engines|FAIRING BLOWER TURBOMACHINE|
    GB202005025D0|2020-04-06|2020-05-20|Rolls Royce Plc|Gearboxes for aircraft gas turbine engines|GB2566047B|2017-08-31|2019-12-11|Rolls Royce Plc|Gas turbine engine|
    GB2566045B|2017-08-31|2019-12-11|Rolls Royce Plc|Gas turbine engine|
    GB202005025D0|2020-04-06|2020-05-20|Rolls Royce Plc|Gearboxes for aircraft gas turbine engines|
    GB202005028D0|2020-04-06|2020-05-20|Rolls Royce Plc|Gearboxes for aircraft gas turbine engines|
    法律状态:
    2018-04-23| PLFP| Fee payment|Year of fee payment: 2 |
    2018-11-09| PLSC| Publication of the preliminary search report|Effective date: 20181109 |
    2019-04-19| PLFP| Fee payment|Year of fee payment: 3 |
    2020-04-22| PLFP| Fee payment|Year of fee payment: 4 |
    2021-04-21| PLFP| Fee payment|Year of fee payment: 5 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1753856|2017-05-02|
    FR1753856A|FR3065994B1|2017-05-02|2017-05-02|BLOWER ROTOR TURBOMACHINE AND REDUCER DRIVING A LOW PRESSURE COMPRESSOR SHAFT|FR1753856A| FR3065994B1|2017-05-02|2017-05-02|BLOWER ROTOR TURBOMACHINE AND REDUCER DRIVING A LOW PRESSURE COMPRESSOR SHAFT|
    PCT/FR2018/000107| WO2018202962A1|2017-05-02|2018-05-02|Turbomachine with fan rotor and reduction gearbox driving a low-pressure decompressor shaft|
    US16/610,095| US11268450B2|2017-05-02|2018-05-02|Turbomachine with fan rotor and reduction gearbox driving a low-pressure decompressor shaft|
    EP18726514.5A| EP3619417A1|2017-05-02|2018-05-02|Turbomachine with fan rotor and reduction gearbox driving a low-pressure decompressor shaft|
    CN201880033553.3A| CN110651112A|2017-05-02|2018-05-02|Turbomachine having a fan rotor and a reduction gearbox driving a shaft of a low-pressure compressor|
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