• 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 for an aircraft, comprising, from upstream to downstream in the direction of a main flow in the turbomachine, at least one gas generator (2a, 2b) configured to conduct a primary flow formed of the all the gases of said generators and a power turbine (3) fed by said central vein (4), said power turbine (3) driving at least one fan (7, 8) and said at least two gas generators (2a, 2b) being arranged so as to leave a free volume around an axis (XX) of the central duct characterized in that it comprises a free turbine rotor (15), installed in the central duct (4), decoupled from the power turbine (3) and configured to convert a portion of the energy of the primary flow into mechanical power on an auxiliary rotary shaft (18), configured to be coupled to at least one auxiliary equipment and out of the central vein in the direction of of said volume libr e. The invention also relates to an aircraft using this turbomachine, in particular at the rear end of its fuselage (1).
    公开号:FR3039206A1
    申请号:FR1556957
    申请日:2015-07-22
    公开日:2017-01-27
    发明作者:Antoine Jean-Philippe Beaujard；Nicolas Jerome Jean Tantot
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    Turbomachine for an aircraft having a free turbine in the primary flow
    Field of the invention and state of the art:
    The present invention relates in particular to the field of turbomachines for aircraft such as airplanes, in particular civil aircraft, having counter-propulsion propulsion blowers placed downstream of the gas generating part driving the power turbine portion coupled to said blowers. It relates more particularly to means integrated in such a turbomachine to supply mechanical power to auxiliary equipment, in particular an electricity generator.
    The type of turbomachine concerned is found, for example, in an aircraft architecture proposed in the patent application FR-A1-2 997 681. In this case, the turbomachine is integrated in the extension of the fuselage downstream thereof , in order to reduce noise and fuel consumption of the aircraft by limiting the aerodynamic drag by absorption of the boundary layer.
    In such an architecture, an aircraft is propelled by a turbomachine with contra-rotating fans careened, the turbomachine being integrated into the rear of the fuselage of the aircraft. Generally, the turbomachine comprises at least two gas generators which feed a power turbine having two counter-rotating rotors for driving two blowers disposed downstream of the gas generators. The gas generators have separate side air intakes to supply each of them.
    Downstream of the gas generators, the blowers are arranged in the extension of the fuselage of the aircraft and generally fed by an annular ring connected thereto, so as to absorb at least a portion of the boundary layer formed around the fuselage. The diameter of the blowers is of the order of that of the fuselage in its largest section. The speed of rotation of the blowers is generally lower than for conventional turbomachines, especially so that the speed at the head of the blade is subsonic.
    As for other types of apparatus, there is a need to use the turbomachine to drive auxiliary equipment, such as an electricity generator for the aircraft. In particular, when the power drawn is relatively large, it is known to couple the equipment to be driven to a power turbine driving the fan rotors, rather than the gas generator so as not to disturb the operation of the latter. The document FR-A1-2941492 shows in particular an integrated architecture, where an electric generator whose rotor is centered on the axis of the turbomachine and is rotated by one of the power turbines.
    The solutions mentioned above may pose problems in the case of an integrated turbomachine at the rear tip of a fuselage, as described, since the power turbine is at the rear end of the aircraft. . There are in this case few fixed housings that can serve as a support for an intermediate equipment, behind the engine or at the nacelle surrounding the blowers. Moreover, this thruster architecture frees up space in the rear tip of the fuselage which it is desirable to use.
    The object of the present invention is to provide a solution for driving auxiliary equipment, in particular a current generator, in the case of the particular architecture of a thruster with a power turbine arranged in the extension of a fuselage. or, more generally, after a vein bringing the gases of the primary flow to the turbine.
    SUMMARY OF THE INVENTION For this purpose, the invention relates to a turbomachine for an aircraft, comprising, from upstream to downstream in the direction of a main flow in the turbomachine, at least two gas generators, a configured central vein for conducting a primary flow formed by all the gases of said generators and a power turbine supplied by said central vein, said power turbine comprising at least one rotor driving at least one fan and said at least two gas generators being arranged to leave a free volume around an axis of the central vein characterized in that it comprises a free turbine rotor, installed in the central duct, decoupled from the power turbine and configured to transform part of the energy primary flow in mechanical power on an auxiliary rotary shaft, configured to be coupled to at least one auxiliary equipment and leaving the a central vein in the direction of said free volume.
    The notion of auxiliary equipment refers, in this document, to equipment not directly involved in the production of a thrust by the turbomachine but which, by performing functions, such as power generation or circulation of fluids , participates in the operation of either the turbomachine directly or the aircraft more generally. The installation of the intermediate rotor in the central duct, uncoupled from the power turbine, makes it possible to operate this intermediate rotor as soon as the gas generator (s) operate, independently of the operating conditions of the propulsion stage, by the power turbine. In addition, the positioning of the intermediate rotor upstream of the power turbine can make it possible to return the power recovered on the auxiliary shaft to the volume left free upstream of the power turbine in that the gas generators are spaced apart. This is particularly interesting for an architecture as described in the introduction, for which it is difficult to install equipment downstream of the turbine or in the nacelle.
    Preferably, said axis of the central vein is aligned with that of the power turbine. This makes it possible to position the auxiliary equipment in the free volume which is upstream of the turbine. Furthermore, in the case of a turbomachine installed in the extension of the rear fuselage of the aircraft, it allows the use of the aircraft structure to install this auxiliary equipment in a fixed housing, which simplifies the installation.
    Advantageously, the free turbine rotor is arranged to improve the homogeneity of the primary flow before entering the power turbine. Indeed, unlike other turbine rotors, the free turbine rotor is here designed to draw from the primary flow the power just necessary for the operation of the auxiliary equipment and leave the main part of the power for the power turbine which is downstream. This leaves the means to design the blades of the free turbine rotor so as to take advantage of their rotation in the primary flow to mix it. This is particularly interesting in the case where the central vein collects the gases of several gas generators which must be homogenized as much as possible before entering the power turbine.
    First rectifying means may be installed in the primary flow, upstream of the free turbine rotor and arranged to improve the efficiency of said free turbine rotor. Said first rectifying means may be formed by a specific arrangement of exhaust casings of the gas generators, at the inlet of the primary vein, or by a crown of radial vanes in the central vein.
    In a conjugated or independent manner, second rectifier means of the primary flow may also be arranged downstream of the free turbine rotor, so as to improve the efficiency of the free turbine rotor. These second means may be formed by a specific arrangement of a distributor at the input of the power turbine.
    Advantageously, said turbomachine comprises said at least one auxiliary equipment, for example an electric current generator, disposed in said free volume left by the gas generators around the axis.
    Preferably, a rotor of said auxiliary equipment is centered on an axis of the power turbine.
    Even more preferably, the turbomachine is configured such that the rotor of the electric current generator rotates at the same speed as the turbine intermediate rotor. The invention also relates to an aircraft propelled by a turbomachine as described above, the central vein of said turbomachine being integrated in the rear of a fuselage of the aircraft and the power turbine being in the extension thereof, said power turbine having two counter-rotating turbine rotors for driving two counter-rotating fans and arranged outside periphery of the power turbine.
    Advantageously, the part of the central vein in which the free turbine rotor is disposed is surrounded by the fuselage of the aircraft.
    Also preferably, the auxiliary equipment driven by the free turbine rotor is inside the fuselage.
    Advantageously, in said aircraft, the turbomachine comprises separate air inlets for supplying each gas generator.
    Brief description of the figures:
    The present invention will be better understood and other details, characteristics and advantages of the present invention will appear more clearly on reading the description of a nonlimiting example which follows, with reference to the appended drawings in which: FIG. a schematic view in longitudinal section of the rear part of an aircraft equipped with a turbomachine according to one embodiment of the invention; - Figure 2 shows a schematic longitudinal sectional view of the rear portion of an aircraft equipped with a turbomachine according to one or more variants of the embodiment of Figure 1; - Figure 3 schematically shows a partial circumferential section of a blade ring in a turbine rotor used by the invention; and FIG. 4 schematically shows an enlargement of part A of FIGS. 1 and 2, in a particular example of equipment concerned by the invention.
    DESCRIPTION OF AN EMBODIMENT The invention applies in particular to an aircraft such as an airplane comprising a turbomachine of the type shown in FIG.
    As shown in FIG. 1, the turbomachine is centered on the longitudinal axis XX of the fuselage 1 of the aircraft. This turbomachine comprises, from upstream to downstream, in the gas flow direction, at least two separate gas generators 2a, 2b simultaneously supplying a single power turbine 3. The turbomachine is installed at the downstream end of the fuselage 1 of the aircraft.
    In this document, the axial and radial designations refer to the axis XX of the fuselage and the turbomachine. Similarly, upstream and downstream terms refer to the direction of the main flow along this axis.
    In a manner known per se, each gas generator 2a, 2b comprises at least one compressor, a combustion chamber and at least one turbine (not shown in the figures).
    Each gas generator 2a, 2b is housed inside a primary flow vein 3a, 3b. Separate air inlets 4a, 4b are provided for these veins 3a, 3b to supply each gas generator 2a, 2b. In the example shown, the air inlets 4a, 4b are connected to the fuselage 1 of the aircraft, upstream of the gas generators 2a, 2b, and their inner wall is directly integrated with the fuselage 1. They thus absorb a portion of the boundary layer formed around the fuselage 1 of the aircraft. In another configuration, not shown, the lateral air inlets feeding each of the gas generators may, on the contrary, be spaced from the fuselage 1 of the aircraft, so as to minimize this phenomenon of absorption of the boundary layer and to facilitate the operation of gas generators. It is also conceivable to use more than two gas generators, for example three to supply the power turbine 3.
    Preferably, the two primary flow veins 3a, 3b of the gas generators 2a, 2b converge on the longitudinal axis XX and form between them an open V upstream, the opening angle is preferably included between 80 ° and 120 °.
    The two primary flow streams 3a, 3b of the gas generators 2a, 2b converge in a central primary stream 4 which feeds the power turbine 3. A mixer (not shown in the figures) is preferably positioned at the level of the zone convergence of the two veins 3a, 3b, housing the gas generators 2a, 2b. This mixer has the function of mixing the gas flows from the two gas generators 2a, 2b to create a single homogeneous gas stream at the outlet of the primary central vein 4.
    The power turbine 3, which is fed by this primary flow output of the central vein 4, is provided with two rotors 5, 6 counter-rotating turbine to drive contrarotatively two fan rotors 7, 8. These turbine rotors 5 , 6 are coaxial and centered on the longitudinal axis XX. They revolve around an inner casing 9 fixed to the structure of the aircraft.
    Here, a first turbine rotor 5 corresponds to vanes connected to a tubular body 5a separating the primary flow vein, in the power turbine 3, from the secondary flow duct, in which the fan rotors 7 are located. 8. The blades and the tubular body 5a of the first rotor 5 are connected to the support bearings of the rotor 5 on the inner casing 9 by support arms 10 which pass through the primary vein upstream of the power turbine 3.
    In the same example, the second rotor 6 corresponds to blades connected to a radially inner wall of the primary stream in the turbine 3 and inserted longitudinally between the vanes of the first rotor 5.
    Downstream of the power turbine 3, the radially inner portion of the second rotor 6 is extended by a central body 11. On the other hand, it is connected by support arms 12 to a ring 13 for supporting the blades of the rotor. In addition, this ring 13 extends the tubular body 5a of the first rotor 5 and has a rearward extension, so as to form, with the central body 11, a primary discharge nozzle, at the outlet of the the power turbine 3.
    In the example shown, a first upstream fan rotor 7 is positioned at the inlet of the power turbine 3. It is connected to the first rotor 5 of the power turbine 3 at the arms 10 which support upstream the cylindrical outer body 5a. This upstream fan rotor 7 therefore rotates at the same speed as the first rotor 5 of the power turbine 3.
    On the same example, the second fan rotor 8, downstream, is positioned at the outlet of the power turbine 3. It is connected to the second rotor 6 of the power turbine 3 at the level of the support ring 13 and arms 12 that support it. This downstream fan rotor 8 therefore rotates at the same speed as the second rotor 6 of the power turbine 3.
    The two blowers 7, 8 are careened by a nacelle 14 fixed to the structure of the aircraft. This nacelle 14 can be fixed, for example, to the vertical tail of the aircraft, not shown in the figures. The blowers have an outer diameter D which is close to the outermost diameter of the fuselage 1 of the aircraft. The air entering the blowers 7, 8 is partly composed of the fuselage boundary layer of the aircraft, the input speed is low compared to conventional turbomachine blowers and the output speed is also lower at identical compression ratio, which improves the propulsive and acoustic performances of these blowers. Moreover, the large outer diameter D of the blowers 7, 8 causes their rotational speed, like that of the rotors 5, 6 of the power turbine 3, will also remain low compared to a conventional turbomachine.
    According to the invention, an intermediate turbine rotor 15, centered on the axis XX, is placed in the central vein 4. The intermediate turbine rotor 15 comprises radial vanes 16 fixed to a central hub 17. The outside diameter of the vanes radial 16 is generally substantially equal to the internal diameter of the central vein 4, leaving a clearance with the walls to allow free rotation.
    The turbine intermediate rotor 15 is designed to operate as a free turbine in the primary flow. For this, as illustrated in Figure 3, the radial vanes 16 are arranged in a ring around the central hub 17 and their axial profile is arranged such that the primary flow F is relaxed in the inter-blade passage. In doing so, the aerodynamic forces on the vanes drive the rotor at a rotational speed ω, in a direction defined by the profile of the vanes 16, as illustrated in FIG.
    Advantageously, the turbine intermediate rotor 15 is designed to deliver a given power as a function of its rotational speed ω, for a range of operating speeds of the gas generators 2a, 2b. This power represents only a small percentage of the energy of the primary flow, the rest of which is used by the power turbine 3.
    The turbine intermediate rotor 15 installed in the central vein 4 is secured to an auxiliary rotary shaft 18 which holds it in the central vein, rotating about the axis XX. Here, the auxiliary shaft 18 has no connection with either of the rotors 5, 6 of the power turbine 3. Its rear end can be located substantially at the turbine intermediate rotor 15. By against, the auxiliary shaft 18 extends forwardly through the central vein 4 at the confluence between the gas inflow from the two gas generators 2a, 2b. It is therefore preferably maintained at the structure of the aircraft or at a casing of the turbomachine, by bearings at this level.
    As can be seen in Figure 1, the gas generators 2a, 2b release a volume in front of the central vein 4 because they deviate to reach the air inlets 4a, 4b which are placed laterally on the fuselage 1. This volume is used here, at least in part, to install equipment 19 able to absorb the power supplied by turbine intermediate rotor 15.
    Although it is sought to recover only a small portion of the power of the primary flow, several variants are possible here to improve the efficiency of the intermediate turbine rotor 15 and to minimize its disturbances on the power turbine 3.
    According to a first variant, the exhaust casings 20 of the gas generators 2a, 2b are designed to orient the gases that emerge in the manner of a turbine stator at the level of their entry into the central duct 4. Exhaust housings 20 may for example comprise grids shaped in particular to redirect the gases along the axis XX in a homogeneous manner. In this way, the vanes 17 of the turbine intermediate rotor 15 can be optimized to work in a homogeneous flow.
    According to a second variant, a ring 21 of stator vanes is installed upstream of the intermediate turbine rotor 15. The vanes are arranged therein to straighten the primary flow in a manner similar to that of a distributor in front of a rotor stage in a turbine classic. This variant makes it possible to minimize the energy loss due to the rotation of the primary flow at the passage of the intermediate turbine rotor 15.
    According to a third variant, rectifying means are placed downstream of the intermediate turbine rotor 15, in particular with a view to providing a flow adapted to the rotors 5, 6 of the power turbine 3. In the example shown in FIG. means comprise fixed blades 22 installed in the divergent part of the distributor of the rotors 5, 6 of the power turbine 3.
    It is also possible to combine these different variants.
    Note also that the free turbine rotor being designed to draw a small part of the power of the primary flow, that just needed for the operation of the auxiliary equipment, it leaves the way to design, by methods known to man of the craft, the blades of the free turbine rotor so as to take advantage of their rotation in the primary flow to mix and promote the homogenization of the gases of the generators which are collected in the central vein, before entering the turbine of power.
    According to another aspect of the invention, with reference to FIG. 4, the equipment 19 absorbing the power of the auxiliary shaft may be an electric current generator.
    As can be seen in FIG. 4, the auxiliary shaft 18 is preferably hollow. Advantageously, it is also open at its downstream end. This allows the passage of easements 23 downstream of the central vein 4, for example oil lines or control rods of the power turbine 3 and / or blowers 7, 8. The upstream end of the auxiliary shaft 18 here kept free to rotate in a fixed housing 24 by bearing bearings 25, 26.
    Advantageously, the electric current generator 19 is formed in the fixed housing 24 by a rotor 27, integral with the auxiliary shaft 18, and by a stator 28, fixed to walls of the fixed housing 24 and surrounding the stator 27. Contacts sliding, not shown, allow to send the current created by the rotation of the rotor 27 to an electrical circuit of the aircraft.
    In a preferred embodiment, the turbine intermediate rotor 15 and the current generator 19 are designed so that the turbine intermediate rotor 15 rotates at a speed that is compatible with the operating characteristics of the generator 19. This avoids adding to the weight of the generator. together with a gearbox between the electric power generator 19 and the turbine intermediate rotor 15.
    However, it is possible to envisage a gear transmission between the turbine intermediate rotor 15 and the generator 19. Additional gears can also be envisaged to drive other auxiliary equipment. The invention has been presented preferably in the case of a turbomachine integrated at the rear tip of a fuselage 1 of an aircraft, with two gas generators 2a, 2b placed laterally. In this configuration, the invention advantageously exploits the existence of a free volume inside the fuselage 1 of the aircraft between the gas generators 2a, 2b, the structure of the aircraft being arranged to allow to install equipment in this free volume. However, an intermediate turbine rotor 15 recovering part of the energy of the primary flow before the power turbine 3 can be installed more generally in the case where the turbojet engine comprises a primary central stream 4 between more than two gas generators and the power turbine.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    claims
    1. Turbomachine for an aircraft, comprising, from upstream to downstream in the direction of a main flow in the turbomachine, at least two gas generators (2a, 2b), a central vein (4) configured to conduct a primary flow formed of all the gases of said generators and a power turbine (3) fed by said central vein (4), said power turbine (3) driving at least one fan (7, 8) and said at least two generators of gas (2a, 2b) being arranged to leave a free volume around an axis (XX) of the central vein characterized in that it comprises a free turbine rotor (15) installed in the central vein (4). ), decoupled from the power turbine (3) and configured to convert a portion of the primary flow energy into mechanical power on an auxiliary rotary shaft (18), configured to be coupled to at least one auxiliary equipment and output from the central vein in direction said free volume.
    [2" id="c-fr-0002]
    2. A turbomachine according to claim 1, wherein said axis (XX) of the central vein is aligned with that of the power turbine (3).
    [3" id="c-fr-0003]
    3. Turbomachine according to one of the preceding claims, wherein the free turbine rotor (15) is arranged to improve the homogeneity of the primary flow before entering the power turbine (3).
    [4" id="c-fr-0004]
    4. A turbomachine according to one of the preceding claims, wherein first rectifying means (20, 21) are installed in the primary flow, upstream of the free turbine rotor (15) and arranged to improve the efficiency of said rotor rotor. free turbine (15).
    [5" id="c-fr-0005]
    5. Turbomachine according to one of the preceding claims, comprising said at least one auxiliary equipment, for example a generator (19) of electric current, disposed in said free volume left by the gas generators around the axis (XX).
    [6" id="c-fr-0006]
    6. Turbomachine according to the preceding claim, wherein a rotor (27) of said auxiliary equipment (19) is centered on an axis (XX) of the power turbine (3).
    [7" id="c-fr-0007]
    7. Aircraft propelled by a turbomachine according to one of claims 1 to 6, the central vein (4) of said turbomachine being integrated in the rear of a fuselage (1) of the aircraft and the power turbine (3 ) being in the extension thereof, said power turbine (3) comprising two contrarotating turbine rotors (5, 6) for driving two counter-rotating fans (7, 8) and arranged at the outer periphery of the power turbine (3). ).
    [8" id="c-fr-0008]
    8. Aircraft according to the preceding claim, wherein the portion of the central vein (4) in which is disposed the free turbine rotor (15) is surrounded by the fuselage (1) of the aircraft.
    [9" id="c-fr-0009]
    9. Aircraft according to the preceding claim, wherein the auxiliary equipment (19) driven by the free turbine rotor (15) is inside the fuselage.
    [10" id="c-fr-0010]
    10. Aircraft according to the preceding claim, wherein the turbomachine comprises separate air inlets (4a, 4b) for supplying each gas generator (2a, 2b).
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    同族专利:
    公开号 | 公开日
    FR3039206B1|2017-07-21|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    FR1385886A|1964-03-05|1965-01-15|English Electric Co Ltd|Turbo-generator for the production of alternating or direct current|
    FR2941492A1|2009-01-23|2010-07-30|Snecma|POWER TURBINE TURBINE ENGINE COMPRISING AN ELECTRONIC POWER GENERATOR CENTERED ON THE AXIS OF TURBOMACHINE|
    FR2997681A1|2012-11-08|2014-05-09|Snecma|PLANE PROPELLED BY A TURBOREACTOR WITH CONTRAROTATIVE BLOWERS|FR3064670A1|2017-03-28|2018-10-05|Safran Power Units|ELECTRIC PRODUCTION INSTALLATION WITH FREE TURBINE|
    FR3072947A1|2017-10-30|2019-05-03|Airbus Operations|AIRCRAFT COMPRISING AT LEAST ONE ENGINE ASSEMBLY CONNECTED TO THE FUSELAGE OF THE AIRCRAFT BY TWO PUSHED CONNECTING LINKS POSITIONED AT LEAST PARTIALLY IN AN AIR INLET OF THE ENGINE ASSEMBLY|
    WO2020008147A1|2018-07-04|2020-01-09|Safran Aircraft Engines|Aircraft propulsion system and aircraft powered by such a propulsion system built into the rear of an aircraft fuselage|
    法律状态:
    2016-08-04| PLFP| Fee payment|Year of fee payment: 2 |
    2017-01-27| PLSC| Search report ready|Effective date: 20170127 |
    2017-05-02| PLFP| Fee payment|Year of fee payment: 3 |
    2018-06-21| PLFP| Fee payment|Year of fee payment: 4 |
    2018-09-14| CD| Change of name or company name|Owner name: SAFRAN AIRCRAFT ENGINES, FR Effective date: 20180809 |
    2019-06-21| PLFP| Fee payment|Year of fee payment: 5 |
    2020-06-23| PLFP| Fee payment|Year of fee payment: 6 |
    2021-06-23| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1556957A|FR3039206B1|2015-07-22|2015-07-22|TURBOMACHINE FOR AIRCRAFT COMPRISING A FREE TURBINE IN THE PRIMARY FLOW|FR1556957A| FR3039206B1|2015-07-22|2015-07-22|TURBOMACHINE FOR AIRCRAFT COMPRISING A FREE TURBINE IN THE PRIMARY FLOW|
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