• 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 present invention relates to an aircraft comprising a fuselage (1) and a thruster downstream of the fuselage, the thruster comprising a turbine (7) of power inside a primary flow vein and at least one fan (9). within a secondary flow stream and driven mechanically by the power turbine, the primary flow stream of the power turbine and the secondary flow stream of the blower being concentric, the power turbine being supplied with gas from two generators (3, 5) of gas turbine gas by two supply channels, characterized in that said gas turbine generators are parallel axes to that of the fuselage, their entry handle d air (3A; 5A) being spaced from the fuselage and the supply channels 3C; 5C) each have a trap (3v; 5v) for controlling the flow between a guiding position of the gas flow towards the power turbine and a position of ejection of the gases in the atmosphere bypassing the power turbine.
    公开号:FR3039133A1
    申请号:FR1556952
    申请日:2015-07-22
    公开日:2017-01-27
    发明作者:Nicolas Jerome Jean Tantot
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    Field of the invention
    The present invention relates to the aeronautical field in which the aircraft are propelled by a set of fan rotors arranged at the rear in the extension of the fuselage. The fan rotors are driven by free turbines, counter-rotating, powered by gas generators formed from turbojets.
    State of the art it has been proposed in patent application FR-A-2 997 681, a new aircraft architecture to reduce noise and fuel consumption of the aircraft by limiting aerodynamic drag.
    In such an architecture, an aircraft is propelled by a propulsion system with counter-rotating fans, integrated in the rear of the fuselage of the aircraft, in the extension thereof. The propulsion system comprising at least two gas generators which feed a power turbine having two counter-rotating rotors for driving two fan rotors, the blowers being arranged downstream of the gas generators, and partially fed by the boundary layer developed in the vicinity of the fuselage of the plane. Gas generators are gas turbine engines incorporated in the fuselage with separate air intakes that each feed a gas generator. Said air intakes are arranged laterally with respect to the fuselage of the aircraft, and absorb at least a portion of the boundary layer formed around the fuselage. The diameter of the nacelle enveloping the fan rotors is, in this embodiment, substantially equal to that of the largest section of the fuselage of the aircraft. This nacelle integrates the power turbine.
    In such a solution, in case of failure of the part of the propulsion system composed of the counter-rotating turbine and the counter-rotating blowers, the maintenance of the ability to generate thrust may not be ensured. For example, in case of destruction of the vanes of the counter-rotating turbine, a total obstruction of the hot internal flow by the debris can occur, it would result the total loss of thrust and a probability of high pumping gas generators in because of the sudden change of debiting section.
    Moreover, the installation mode of the gas generators fully integrated into the fuselage assumes an arrangement of the air inlets to not feed the generators with the boundary layer developed on the aircraft fuselage; the latter having a speed substantially less than the flight speed, is detrimental to the thermal efficiency of the gas generators.
    The object of the present invention is to provide a solution adapted to the type of aircraft and propeller architecture that has just been described so that the assembly operates optimally independently of the different flight conditions.
    Thus, more particularly, an objective is to feed the gas generators without ingestion of fuselage boundary layer or flow distortion, to maximize the thermal efficiency, it is also to maintain the ability to power the thruster with the maximum of air from the boundary layer to maximize the propulsive efficiency.
    The present invention also aims to maintain a thrust capacity in case of major failure of the propellant module.
    The present invention also aims, in the event of a failure of the thruster, to ensure the segregation of the flows between the functional parts of the propulsion system and the parts that are no longer functional.
    The present invention also aims to implement an installation of gas generators compatible with the certification requirements, particularly vis-à-vis the risks of "cross-debris". By this term denotes the damage of one of the gas generators by the debris from the other generator.
    Presentation of the invention
    These objectives are achieved with an aircraft having a fuselage and a thruster downstream of the fuselage, the thruster comprising a power turbine inside a primary flow vein and at least one blower inside a vein of secondary flow and driven mechanically by the power turbine, the primary flow stream and the secondary flow stream being concentric, the power turbine being supplied with gas from two gas turbine gas generators by two supply channels, said aircraft being characterized by the fact that said gas turbine generators are parallel axes to that of the fuselage, their air inlet sleeve being spaced from the fuselage and the supply channels each have a control flap of mobile flow between a guiding position of the gas flow towards the power turbine and a position of ejection of the gases coming from the gas generator into the atmosphere era bypassing the power turbine.
    The solution of the invention differs from that presented in patent EP 1267064 in the name of the present applicant which relates to a variable configuration engine operating alternately nominally subsonic and supersonic, where the invention aims to provide a palliative device in the event of failure of the main thruster located in the rear tip of the aircraft; Moreover, the architecture of a supersonic aircraft is not primarily concerned by a problem of optimizing the efficiency of the gas generators, vis-à-vis ingestion of the ingested boundary layer developed on the fuselage. The invention on the contrary ensures the positioning of the generators far enough away from the fuselage so that their inlet hoses are spaced from the fuselage and they do not ingest this boundary layer, penalizing in terms of the thermal efficiency of the aircraft. engine. In addition, the arrangement proposed by the invention on either side of the fuselage limits the risk of "cross debris".
    Preferably, at least two gas generators spaced from any wing of the aircraft.
    Even more preferably, the aircraft comprises two generators hidden relative to each other by the fuselage of the aircraft.
    Here the risk of "cross debris" is further reduced due to the presence of the fuselage between the two gas generators.
    Advantageously, the supply channels comprise means for ejecting gases into the atmosphere which form wall elements of said supply channels when the hatches are positioned to guide the gas flow to the power turbine.
    Indeed, the propulsion of the aircraft is optimized to operate the thruster downstream of the gas generators. Their integration in the wall of the supply ducts makes it possible to minimize the disturbances of the external flow during this operation.
    More particularly the control flap of the air flow from each gas generator comprises a wall element movable between a closed position of the feed channel and ejection opening of the flow parallel to the axis of the fuselage , and a guide position of the flow in the feed channel.
    According to another feature, the two power supply channels converge upstream of the power turbine into a single power supply channel.
    Advantageously, the generators are single-flow turbojets and more particularly the turbojets are double-body.
    According to an advantageous embodiment, the thruster comprises a power turbine formed of two counter-rotating rotors each driving a fan rotor.
    In particular, the ejection channel from the turbojets is convergent. Thus in the invention, the hot gas ejection in the case of the main propulsion failure (considered as a backup mode) is performed through a thrust reverser type gate, whose passage section is made slightly convergent, thus acting as a simplified nozzle and generating thrust in the axis of the aircraft.
    Presentation of the Figures The invention will be better understood, and other objects, details, features and advantages thereof will appear more clearly on reading the following detailed explanatory description of an embodiment of the invention. given by way of purely illustrative and non-limiting example, with reference to the appended schematic drawing. On these drawings:
    Figure 1 is a schematic representation of a propulsion assembly according to the invention; the upper half shows the hatch in the normal operating position of the engine, the lower half shows the hatch in the shut-off position of the turbine feed channel.
    FIG. 2 shows the installation of the propulsion unit of FIG. 1 on an aircraft according to the invention.
    DETAILED DESCRIPTION OF AN EMBODIMENT OF THE INVENTION
    With reference to FIGS. 1 and 2, the propulsion unit is mounted at the rear of the aircraft, on the rear part of the fuselage 1. On this fuselage, the two gas generators 3 and 5 are mounted by means of pylons 3p and 5p respectively. These two pylons allow to reserve sufficient space between the wall of the fuselage and the air intake sleeve of each of the engines, 3a and 5a thus avoiding that the air of the boundary layer formed along the fuselage is directed towards the air intake hoses of the gas generators, while feeding the blowers 91 and 93 of the thruster.
    As can be seen in FIG. 2, the positioning of the pylons 3p and 5p, here laterally at the rear of the fuselage 1, moves the generators 3 and 5 away from the wings 12, situated substantially in the middle part of the fuselage, and tail tail, formed here of a vertical drift 13 supporting two fins 14. More specifically, the gas generators are at the height of the fuselage arranged laterally and the fins are vertically away from the fuselage. In addition, the two generators 3, 5 are here installed upstream of the tail of the fuselage 1, in a region where the fuselage diameter allows the latter to hide each generator with respect to the other and the fins are at a distance from the gas turbine turbines to prevent a turbine disk burst from causing damage to the other gas generator or the fins.
    The gas generators are, in the example illustrated here, single-flow turbofan engines with double bodies. they thus comprise a low pressure body formed of a rotor with a compressor 31; 51 and a turbine 39; 59, and a high pressure rotor formed of a compressor 33; 53 and a turbine 37; 57. The compressors supply air to a combustion chamber 35; 55 whose gases produced drive the high pressure and low pressure turbines successively. The gases are guided to the exhaust ducts 3t and 5T immediately downstream of the turbines. These channels are extended by conduits 3C and 5C which converge to an inlet conduit 7c of the thruster located downstream of the fuselage.
    The thruster comprises a turbine 7 formed of two coaxial and counter-rotating turbine rotors, 71 and 73, in the axis of the fuselage. The gas stream driving the counter-rotating turbine constitutes the primary flow vein. Each turbine rotor is mechanically secured to an outer concentric fan rotor 91 and 93 respectively. The counter-rotating fan 9 rotates inside the fan casing 10 which is connected by arms 11 to the fuselage 1 and which defines the vein of the fan. secondary stream.
    The traps 3v and 5v are arranged downstream of the exhaust ducts 3t and 5t of the two gas generators 3 and 5. They can pivot about an axis located downstream with respect to the respective exhaust channel. These traps form elements of the guiding wall of the gas stream from the exhaust channel 3T and 5t.
    In normal operation, the gases coming from the exhaust channels are guided in the conduits 3C and 5C for supplying the thruster. The two flows converge towards the input 7C of the thruster and constitute the primary flow which rotates the counter-rotating turbine 7 before being ejected into the atmosphere through the primary flow nozzle 15. The two turbine rotors 71 and 73 rotate each of the two rotors 91 93 of the counter-rotating fan 9. These aspirate the outside air in the volume defined by the fan casing and the hull of the primary vein. The air passing through the fan constitutes the secondary flow.
    Furthermore, in normal operation, each hatch 3v, 5v is positioned to integrate in the extension of the wall, between the exhaust channel 3t, 5t and the intake channel 3C, 5C. In this way, the flow outside the aircraft and the generators is the least disturbed possible.
    When a failure occurs on the thruster with a risk of clogging the vein of the primary flow, which would cause a total loss of propulsion, it is possible with the arrangement of the invention to ensure the propulsion of the aircraft. directly by the gases from the generators 3 and 5. The traps 3v and 5v are pi voted around their axis of rotation so as to obstruct the channels to the ducts 3C and 5C supply of the propellant. By pivoting, the hatches 3v and 5v discover a part, 3D; 5d, of the wall in the axis of the associated exhaust channel 3T and 5T, thus forming a means for exhausting gases from each generator to the atmosphere. Thus the continuity of the propulsion is assured.
    权利要求:
    Claims (10)
    [1" id="c-fr-0001]
    1. Aircraft comprising a fuselage (1) and a thruster downstream of the fuselage, the thruster comprising a turbine (7) of power inside a primary flow vein and at least one blower (9) inside a secondary flow stream and driven mechanically by the power turbine, the primary flow stream of the power turbine and the secondary flow stream of the fan being concentric, the power turbine being supplied with gas from minus two generators (3, 5) of gas turbine gas by at least two supply channels, characterized in that said gas turbine generators are of axes parallel to that of the fuselage, spaced from the fuselage, and the supply channels (3C; 5C) each have a trap (3V; 5v) for controlling the flow between a guiding position of the gas flow towards the power turbine and a position of ejection of the gases into the atmosphere in bypassing a power turbine.
    [2" id="c-fr-0002]
    2. Aircraft according to claim 1, comprising at least two gas generators spaced from any wing (12, 13, 14) of the aircraft.
    [3" id="c-fr-0003]
    3. Aircraft according to one of claims 1 and 2, comprising two generators masked relative to each other by the fuselage (1) of the aircraft.
    [4" id="c-fr-0004]
    4. Aircraft according to one of the preceding claims, wherein the supply channels (3c; 5C) comprise means for ejecting gases into the atmosphere which form wall elements of said supply channels when traps (3v, 5v) are positioned to guide the gas flow to the power turbine (7).
    [5" id="c-fr-0005]
    5. Aircraft according to the preceding claim, the flap (3v; 5v) flow control comprises a movable wall element between a closed position of the feed channel and ejection opening of the flow parallel to the fuselage axis and a guide position of the flow in the feed channel
    [6" id="c-fr-0006]
    6. Aircraft according to one of the preceding claims, wherein the two supply channels (3C; 5C) converge upstream of the power turbine in a single channel (7c) supply.
    [7" id="c-fr-0007]
    7. Aircraft according to one of the preceding claims, whose generators are single-flow turbojets,
    [8" id="c-fr-0008]
    8. Aircraft according to the preceding claim, the turbojet engines are double body.
    [9" id="c-fr-0009]
    9. Aircraft according to one of the preceding claims, wherein the thruster comprises a power turbine (7) formed of two rotors (71; 73) counter-rotating, each driving a rotor (91; 93) blower.
    [10" id="c-fr-0010]
    10. Aircraft according to the preceding claim, the ejection channel from turbojets is convergent.
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    同族专利:
    公开号 | 公开日
    EP3325345A1|2018-05-30|
    CN107848629A|2018-03-27|
    US10829232B2|2020-11-10|
    US20180208322A1|2018-07-26|
    EP3325345B1|2020-01-15|
    WO2017013363A1|2017-01-26|
    CN107848629B|2021-06-08|
    FR3039133B1|2020-01-17|
    引用文献:
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    US3099425A|1960-12-16|1963-07-30|Hamburger Flugzeugbau Gmbh|Jet propulsion system|
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    US20150247455A1|2013-08-15|2015-09-03|United Technologies Corporation|External core gas turbine engine assembly|US10883424B2|2016-07-19|2021-01-05|Pratt & Whitney Canada Corp.|Multi-spool gas turbine engine architecture|
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    US11035293B2|2016-09-15|2021-06-15|Pratt & Whitney Canada Corp.|Reverse flow gas turbine engine with offset RGB|
    US10465611B2|2016-09-15|2019-11-05|Pratt & Whitney Canada Corp.|Reverse flow multi-spool gas turbine engine with aft-end accessory gearbox drivingly connected to both high pressure spool and low pressure spool|
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    US10823056B2|2016-12-07|2020-11-03|Raytheon Technologies Corporation|Boundary layer excitation aft fan gas turbine engine|
    US10808624B2|2017-02-09|2020-10-20|Pratt & Whitney Canada Corp.|Turbine rotor with low over-speed requirements|
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    CN111284714A|2019-12-30|2020-06-16|中电科芜湖钻石飞机制造有限公司|Commuting type electric airplane system|
    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
    FR1556952A|FR3039133B1|2015-07-22|2015-07-22|AIRCRAFT WITH A PROPELLANT ASSEMBLY INCLUDING A BLOWER AT THE BACK OF THE FUSELAGE|
    FR1556952|2015-07-22|FR1556952A| FR3039133B1|2015-07-22|2015-07-22|AIRCRAFT WITH A PROPELLANT ASSEMBLY INCLUDING A BLOWER AT THE BACK OF THE FUSELAGE|
    CN201680045187.4A| CN107848629B|2015-07-22|2016-07-21|Aircraft comprising a propulsion assembly with a fan on the rear of the fuselage|
    PCT/FR2016/051885| WO2017013363A1|2015-07-22|2016-07-21|Aircraft comprising a propulsion assembly including a fan on the rear of the fuselage|
    EP16757290.8A| EP3325345B1|2015-07-22|2016-07-21|Aircraft comprising a propulsion assembly including a fan on the rear of the fuselage|
    US15/745,770| US10829232B2|2015-07-22|2016-07-21|Aircraft comprising a propulsion assembly including a fan on the rear of the fuselage|
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