• 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 an aircraft comprising a fuselage, flight control surfaces, and a turbojet engine (20) integrated at the rear of said fuselage in the extension thereof, the turbojet engine (12) comprising two gas generators (22) which supply, by a common central duct (30), a power turbine (32) comprising two counter-rotating rotors (34, 36) respectively driving two coaxial and counter-rotating upstream (38) and downstream (40) blowers each comprising a blade crown (42, 44), the set of blowers (38, 40) being integrated in a shroud (46) of the turbojet engine (20) formed at the rear of the fuselage (12), characterized in that at least said shroud ( 46) is arranged axially behind the flight control surfaces and comprises an upstream portion (50) surrounding the upstream fan (38), configured to be traversed radially by at least one blade fragment (43) (42) of the fan upstream (38), in case of dawn rupture (42) of said souf flante (38) and ejection of said at least fragment (43).
    公开号:FR3054526A1
    申请号:FR1657186
    申请日:2016-07-26
    公开日:2018-02-02
    发明作者:Jerome Jean Tantot Nicolas；Alain Eric Sauve Michael
    申请人:Safran Aircraft Engines SAS；
    IPC主号:
    专利说明:

    © Publication no .: 3,054,526 (to be used only for reproduction orders)
    ©) National registration number: 16 57186 ® FRENCH REPUBLIC
    NATIONAL INSTITUTE OF INDUSTRIAL PROPERTY
    COURBEVOIE © Int Cl 8 : B 64 D 27/14 (2017.01), B 64 D 27/20, 27/00, 33/02
    A1 PATENT APPLICATION
    ©) Date of filing: 07.26.16. © Applicant (s): SAFRAN AIRCRAFT ENGINES - (© Priority: FR. @ Inventor (s): TANTOT NICOLAS, JEROME, JEAN and SAUVE MICHAEL, ALAIN, ERIC. (43) Date of public availability of the request: 02.02.18 Bulletin 18/05. ©) List of documents cited in the report preliminary research: Refer to end of present booklet (© References to other national documents ® Holder (s): SAFRAN AIRCRAFT ENGINES. related: ©) Extension request (s): (© Agent (s): GEVERS & ORES Société anonyme.
    / b AIRCRAFT COMPRISING A TURBOJECTOR INTEGRATED IN THE REAR FUSELAGE COMPRISING A FAIRING ALLOWING THE EJECTION OF BLADES.
    FR 3 054 526 - A1 _ The invention relates to an aircraft comprising a fuselage, flight control surfaces, and a turbojet engine (20) integrated at the rear of said fuselage in the extension thereof, the turbojet engine (12) comprising two gas generators (22) which feed, via a common central stream (30), a power turbine (32) comprising two counter-rotating rotors (34, 36) respectively driving two coaxial and counter-rotating blowers upstream (38) and downstream (40) each comprising a crown of blades (42, 44), the set of blowers (38, 40) being integrated in a fairing (46) of the turbojet engine (20) formed at the rear of the fuselage (12), characterized in that that at least said fairing (46) is arranged axially behind the flight control surfaces and comprises an upstream part (50), surrounding the upstream fan (38), configured to be traversed radially by at least one fragment (43) of blade (42) of the upstream fan (38), in the event of blade failure (42) of said upstream blower (38) and of ejection of said at least fragment (43).

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    i
    Aircraft comprising a turbojet engine integrated into the rear fuselage comprising a fairing allowing the ejection of blades
    Field of the invention and state of the art:
    The present invention relates to the field of aircraft such as airplanes, in particular civil aircraft, powered by a turbofan engine with counter-rotating blowers, integrated in the extension of the fuselage downstream thereof. It relates more particularly to the means for adapting the blowers of such turbojets in this type of architecture to degraded operating situations which must be taken into account for safety reasons.
    It has been proposed, in patent application FR-A1-2 997 681, a new aircraft architecture making it possible to reduce the noise pollution and the fuel consumption of the aircraft by limiting the aerodynamic drag by absorption and re-energization by the propellant system of the boundary layer developed on the fuselage.
    In such an architecture, an aircraft is propelled by a turbofan with keeled contra-rotating blowers, the turbojet being integrated at the rear of the fuselage of the aircraft. The turbojet engine comprises at least two gas generators housed in a rear end of the fuselage, which feed a power turbine having two counter-rotating rotors intended to each drive a corresponding fan arranged downstream of the gas generators. The gas generators are supplied with air by separate lateral air inlets which are arranged laterally relative to the fuselage. Downstream of the gas generators, the respective blowers, said upstream and downstream, are arranged in the extension of the fuselage of the aircraft and are generally supplied so as to absorb at least part of the boundary layer formed around the fuselage. The diameter of the blowers is about that of the fuselage in its largest section. The speed of rotation of the blowers of such a turbojet is generally lower than for conventional turbomachines, in particular so that the speed at the dawn of the blowers is subsonic. The blowers are arranged in a fairing which provides an aerodynamic guiding function, both for the internal flow of air passing through these blowers and for the external air flow which runs along the fairing.
    Such a turbojet engine can be used as the sole engine of the aircraft on which it is mounted, or as the main engine, this turbojet engine then being assisted in its task by two conventional auxiliary turbomachines placed under the wings of the aircraft. In this case, the turbojet engine and the turbomachines are used selectively as a function of the different flight phases of the aircraft, the auxiliary turbomachines being for example only used when an additional thrust is necessary as is the case during the takeoff, and the main turbojet engine being used during cruising flight phases.
    Like a conventional turbojet, a turbojet placed in the extension of the fuselage of an aircraft may experience a failure. For example, it may happen that one of the two blowers loses a dawn. This phenomenon, although rarely caused by the ingestion of a bird due to the positioning of the turbojet engine in the extension of the fuselage of the aircraft, can nevertheless occur following the appearance of defects in the material of a blade. The loss of one or more fan blades risks damaging the entire turbojet engine. In this case it is necessary to limit the rotation of the rotor of the fan concerned to allow the aircraft to return to an airport without major impact on its flight capacity.
    The loss of a fan blade is particularly detrimental when it concerns the fan upstream of the turbojet engine. Indeed, in this case, the loss of one or more blades of the upstream fan can lead said blades to be guided by the fairing up to the downstream fan, which can cause the destruction of the downstream fan and consequently cause a total loss of the propellant capacity of the turbojet engine.
    This possibility is not admissible when the turbojet is the one and only means of propulsion of the aircraft.
    The object of the present invention is to provide a turbojet engine having the characteristics described above, and making it possible to avoid any degradation of the downstream fan due to loss of the blade of the upstream fan. In fact, a loss of blade from the downstream fan is not directly detrimental to the safety of the aircraft, since the blades of the downstream fan or their debris, which are subjected to the flow of air, are not likely to be entrained in the upstream blower, and therefore cannot cause major degradation of the upstream blower.
    A secondary objective of the invention is to maintain satisfactory aerodynamic behavior of the turbojet engine after the loss of a blade from the upstream fan.
    Statement of the invention:
    To this end, the invention proposes an aircraft comprising a turbojet engine integrated at the rear of the fuselage of the aircraft, and equipped with means allowing the radial ejection of any blade debris liable to break in order to prevent this blade or fragments of this blade are entrained towards the downstream fan.
    To this end, the invention proposes an aircraft comprising a fuselage, flight control surfaces, and a turbojet engine integrated at the rear of said fuselage in the extension thereof, the turbojet engine comprising at least two gas generators which supply, by a common central vein, a power turbine comprising two counter-rotating rotors respectively driving two coaxial and counter-rotating blowers upstream and downstream each comprising a crown of blades, all of the blowers being integrated in a fairing of the turbojet engine formed at the rear of the fuselage , characterized in that at least said fairing is arranged axially behind the flight control surfaces and comprises an upstream part, surrounding the upstream fan, configured to be traversed radially by at least one blade fragment of the upstream fan, in case blade failure of said upstream fan and ejection of said at least fragment.
    This configuration advantageously makes it possible to eject any fragment of a blade which would break in a substantially radial direction, in order to prevent it from being entrained towards the downstream fan and not risking damaging it. Furthermore, this configuration also makes it possible to guarantee that no blade fragment is ejected in the direction of the flight control surfaces of the aircraft and does not risk coming into contact with them and damaging them.
    According to another characteristic of the invention, the upstream fairing part is configured to be at least partially torn by the passage of said at least blade fragment during its crossing of said upstream fairing part.
    Furthermore, the invention provides an aircraft comprising palliative means aimed at partially reconstructing the aerodynamic guidance function of the fan shroud, after the latter has been torn apart by the fragments of this blade.
    Thus, according to another characteristic of the invention, the fairing comprises at least one palliative casing which is capable of being deployed on any torn part of the upstream part of the fairing.
    Advantageously, the palliative casing comprises a plurality of angular segments of palliative casing, angularly distributed around the axis of the fairing, each of which is movable between a retracted position in which it is retracted into a housing of the fairing formed outside the upstream part of said fairing, and a deployed position in which it extends in line with the upstream fairing part, substantially in the extension of the fairing, in order to close said torn part,
    The aircraft also comprises a palliative casing system comprising means for controlling the palliative casing which are configured to be activated at least in response to information on the failure of the blade of the upstream fan, and a means for detecting rupture of upstream dawn capable of providing said information.
    Advantageously, the means for controlling the palliative housing include means for deploying each angular segment.
    According to a particular embodiment of the invention, the control system of the palliative housing includes means for identifying the angular position of any torn part of the upstream fairing part capable of providing position information of any torn part, and the palliative case control means are configured to selectively activate the means for deploying each angular sector associated with a torn part of the upstream fairing part in response to said position information.
    The invention also relates to a turbojet engine comprising comprising two counter-rotating rotors respectively driving two coaxial and counter-rotating blowers upstream and downstream each comprising a crown of blades, all of the blowers being integrated in a fairing of the turbojet engine. This turbojet engine is characterized in that said fairing comprises at least one palliative housing which is capable of being deployed over any part of the upstream fairing capable of being traversed radially and at least partially torn by a blade fragment of the upstream fan, for close off the corresponding torn part.
    The invention also relates to a method for controlling a palliative housing system of an aircraft of the type described above, characterized in that it comprises a first step during which the upstream blade failure detection means detects the rupture of at least one blade of the upstream fan, and a second step, determined by the first step, during which the control means control the deployment of the palliative housing.
    The invention also relates to a variant control method of an aircraft palliative casing system of the type described above, characterized in that it comprises a first step during which the upstream blade break detection means detects the rupture of at least one blade of the upstream fan, a second step during which the means for identifying the angular position of any torn part of the upstream fairing part detects the position of any torn part of the part upstream of fairing, and a third step, determined by the first and second steps, during which the control means selectively activate the means for deploying each angular sector associated with a torn part of the upstream fairing part.
    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:
    - Figure 1 is a perspective view of an aircraft produced according to a prior art;
    - Figure 2 is a schematic sectional view of a turbojet engine according to the prior art of the equipment fitted to the aircraft of Figure 1;
    - Figure 3 is a top view of an aircraft produced in accordance with the invention;
    - Figure 4 is a schematic sectional view of a turbojet engine according to first and second embodiments of the invention fitted to the aircraft of Figure 3 during a normal operating phase;
    - Figure 5 is a schematic perspective view with cutaway of a fairing for the turbojet engine of Figure 4;
    - Figure 6 is a schematic sectional view of the turbojet engine of Figure 4 during a loss phase of a blade of the upstream fan;
    - Figure 7A is a schematic sectional view of the turbojet engine according to the first embodiment of the invention during a deployment phase of its palliative housing;
    FIG. 7B is a schematic sectional view of the turbojet engine according to the second embodiment of the invention during a deployment phase of its palliative housing,
    FIG. 8 a flow diagram representative of a method for controlling the palliative housing system according to the first disclosure mode of the invention,
    - Figure nine 9 is a flowchart representative of a method of controlling the palliative system according to the second mode of the invention.
    Description of an embodiment:
    FIG. 1 shows an aircraft 10 produced in accordance with a prior state of the art.
    As shown in FIG. 1, the aircraft 10 comprises a fuselage 12, wings 14 extending from the fuselage 12, and, at a rear end or tail of the fuselage 12, a turbojet engine 20 integrated in the extension of the fuselage 12 , substantially coaxial to an axis XX of the fuselage 12. In front of the turbojet engine 20, the aircraft 10 comprises flight control surfaces comprising a fin 17 carrying a rudder 16 and rudders 18, which extend partially above the turbojet 20.
    In the remainder of this description, the axial and radial designations refer to the axis XX of the fuselage 12 and of the turbojet engine 20. Similarly, the terms upstream and downstream refer to the direction of the main flow along this axis XX.
    The aircraft 10 which has been shown in FIG. 1 comprises a single turbojet engine 20 which forms the main engine of the aircraft. It is noted that there are also, however, aircraft 10 fitted with turbojets of this type and conventional turbomachines additionally fixed under the wings 14, these conventional turbomachines being used as auxiliary engines for certain phases of flight of the aircraft, for example during takeoff of the aircraft 10, in order to provide additional thrust, while the turbojet engine 20 is itself more specifically dedicated to cruising flight phases.
    As illustrated more particularly in FIG. 2, such a turbojet engine 20 comprises two gas generators 22 which are, in this example, supplied with air collected at the surface of the fuselage 12 by means of respective orifices 24 forming inlets of air and via respective conduits 26.
    The orifices 24 thus absorb part of the boundary layer formed around the fuselage 12 of the aircraft 10. In another configuration, not shown, the orifices 24 supplying each of the gas generators 22 can, on the contrary, be separated from the fuselage 12 of the aircraft, so as to minimize this phenomenon of absorption of the boundary layer and to facilitate the operation of the gas generators 22. It is also possible to envisage using more than two gas generators 22.
    In a manner known per se, each gas generator 22 comprises at least one compressor, a combustion chamber and a turbine (not shown in FIG. 2).
    The gas generators 22 supply, via respective conduits 28 delimiting two primary streams 29, a common central stream 30 supplying the combustion gases produced by the gas generators 22 into a power turbine 32. Preferably, the two primary flow streams 29 of the gas generators 22 converge on the longitudinal axis XX and form between them a V open upstream, the opening angle of which is preferably between 80 ° and 120 °.
    The two primary flow streams 29 of the gas generators 22 converge in the central stream 30 which supplies the power turbine 32. A mixer (not shown in the figures) is preferably positioned at the level of a convergence zone 31 of the two primary veins 29. The function of this mixer is to mix the gas flows from the two gas generators 22 to create a single homogeneous gas flow at the outlet of the central vein 30.
    The power turbine 32 comprises two counter-rotating rotors 34, 36 respectively driving two coaxial and counter-rotating blowers upstream 38 and downstream 40. The turbine rotors 34, 36 are coaxial and centered on the longitudinal axis X-X. They revolve around a central casing 35 fixed to the structure of the aircraft. Downstream of the power turbine 32, the radially internal part of the rotor 34 is extended by a central body 37. On the other hand, it is connected, by support arms 39, to a ring 41 for supporting the blades 44 of the downstream fan 40. In addition, this ring 41 has a rearward extension, so as to form, with the central body 37, a primary ejection nozzle, at the outlet of the power turbine 32.
    The upstream fan 38 has a ring of blades 42 and the downstream fan 40 has a ring of blades 44.
    In known manner, all of the blowers 38, 40 are integrated in a fairing 46 of the turbojet engine formed at the rear of the fuselage 12. The fairing 46 is traditionally arranged in the extension of the fuselage 12, as shown in FIGS. 1 and 2.
    This design poses safety problems in the event of rupture of one or more blades of the front fan 38.
    Indeed, conventionally, the fairing 46 is a substantially rigid fairing which, on the one hand, performs aerodynamic functions since it is intended to guide an air flow in a stream of secondary air 48 passing through the blowers 38, 40 , and on the other hand, structural functions, in particular with a view to allowing the retention of blade debris in the event of failure of the upstream fan 38.
    Therefore, in the event of a blade 42 breaking from the upstream fan 38, the fairing 46 is capable of guiding fragments 43 of the broken blade 42 towards the downstream fan 40 into which they can penetrate and cause considerable damage. . This damage can go as far as the destruction of the downstream fan 40, which can lead to a complete loss of propellant capacity of the turbojet engine 20. This risk is not acceptable, in particular when the turbojet engine 20 constitutes the main engine of the aircraft 10, case in which a deficiency of said turbojet engine 20 cannot be remedied by the use of auxiliary turbomachines.
    The invention overcomes this drawback by proposing a new design of the aircraft 10 and in particular of the turbojet 20 capable of preventing the sending of broken blades 42 or blade fragments from the upstream fan 38 in the vein. secondary 48, in order to avoid any risk of deterioration of the downstream fan 40.
    An aircraft 10 produced in accordance with the invention has been shown in FIG. 3.
    The aircraft 10 comprises, as previously, a fuselage 12, wings 14 extending from the fuselage 12, and, at a rear end or tail of the fuselage 12, flight control surfaces comprising a rudder 16 and rudders 18 of depth. The aircraft 10 comprises a main turbojet engine 20 integrated into the rear of its fuselage 12 in the extension thereof. The turbojet engine 20 is arranged substantially coaxially to an axis X-X of the fuselage 12. As before, the turbojet engine 20 could be assisted by auxiliary turbomachines (not shown) placed under the wings 14 of the aircraft 10.
    According to the invention, the fairing 46 of the turbojet engine 20 is arranged axially behind the flight control surfaces 16, 18, that is to say that the upstream fan 38 and the downstream fan 40 are also arranged behind the control surfaces of vol 16, 18.
    More particularly, as illustrated in FIG. 4, a turbojet engine 20 according to the invention comprises, as before, two gas generators 22 each placed in a primary stream 29 which are supplied with air which is for example, and in a nonlimiting manner of the invention, collected on the surface of the fuselage (not shown). For example, each gas generator 22 comprises at least one low pressure compressor BP 22a, a high pressure compressor HP 22b, a combustion chamber
    22c, a high pressure turbine HP 22d and a low pressure turbine BP 22e. The BP compressor 22a and the BP turbine 22e are connected to each other by a low pressure shaft BP 22f, and the HP compressor 22B and the turbine HP 22d are connected to each other by a high pressure shaft HP 22g, which have been shown diagrammatically in FIG. 4. The gas generators 22 supply, via respective conduits 28 delimiting the two primary streams 29, a common central stream 30 supplying the combustion gases produced by the gas generators 22 in the power turbine 32, similar to the previous configuration. The power turbine 32 comprises two counter-rotating rotors 34, 36 respectively driving two blowers 38, 40 which rotate around a central casing 35 fixed to the structure of the aircraft, and which are surrounded by the fairing 46.
    Furthermore, the fairing 46 comprises an upstream part 50, surrounding the upstream blower 38, which is configured to be traversed radially by at least one fragment 43 of the blade of the upstream blower, in the event of the blade 42 breaking of said blower upstream 38 and ejection of said at least fragment 43.
    This configuration advantageously makes it possible to ensure that, in the event of rupture of one or more blades 42 of the upstream fan 38, this or these blades 42, or their fragments 43, are not entrained by the fairing 46 as far as the fan. downstream 40 and therefore do not risk damaging it. This configuration therefore makes it possible to maintain a propulsive capacity of the downstream blower 40 even if the upstream blower 38 is damaged.
    It should be noted that this configuration is specifically intended for the upstream part 50 of the fairing 46 since it is only in the event of a blade 42 of the upstream fan 38 breaking that the turbojet engine 20 can run a significant risk, a rupture a blade 44 of the downstream fan 40 causing a loss of the propulsive capacities of said downstream fan 40 but not likely to cause deterioration of the upstream fan 38, so that the turbojet engine 20 would still be mainly operational.
    The upstream part 50 of the fairing 46 which is capable of being traversed by a fragment 43 of the blade 42 can be produced in any manner known from the state of the art. For example, the upstream part 50 of the fairing 46 could be traversed in a non-destructive manner by a fragment 43 of blade 42 and for this purpose one could consider equipping this upstream part 50 with movable walls capable of disappearing as the blade fragments 42 and then resume their position.
    However, in the preferred embodiment of the invention, the upstream fairing part 50 is configured to be torn at least partially by the passage of said at least fragment 43 of blade 42 during its crossing of said upstream fairing part 50 46, as shown in Figure 6.
    By tear is meant in the broad sense any attack on the integrity of the upstream part 50 of the fairing 46, causing the appearance of an opening through the fairing 46, by analogy with the vocabulary of shipbuilding. For example, a tear in the fairing will in most cases correspond to a piercing of this fairing by a fragment of a blade, or a tearing of part of this fairing by a fragment of a blade.
    The fairing 46 is then a fairing called “soft fairing” which must therefore therefore have characteristics of rigidity sufficient to allow the guiding of the air flow in the secondary vein 48 and which must nevertheless be able to be torn locally by crossing the or fragments 43 of the blade 42 when they are ejected radially due to the rotation of the upstream fan 38. This configuration is made possible by the fact that, unlike the prior art, the fairing 46 is not in this case not intended to provide structural functions.
    This configuration makes it possible to guarantee that, at a minimum, any rupture of a blade 42 of the upstream fan 38 does not cause deterioration of the fin 17, nor deterioration of the downstream fan 40. As a result, the aircraft 10 retains its propulsive and guiding capacities which can allow him to return to an airport.
    However, the invention advantageously proposes palliative means intended to restore, after a tear in the upstream part 50 of the fairing 46, the integrity of this fairing 46.
    To this end, as illustrated in FIGS. 5, 7A and 7B, the fairing comprises at least one palliative housing 52 which is capable of being deployed on any torn part 54 of the upstream part 50 of the fairing 46. When it is torn , as shown in FIGS. 7A and 7B, the fairing 46 includes a torn part 54 on which the palliative housing 52 can be deployed to close said torn part 54 and thus restore the integrity of the fairing 46.
    Any embodiment of such a palliative housing 52 can be used for the proper implementation of the invention. The fairing 46 comprises, in a known manner, an internal wall 56 and an external wall 58. For example, the palliative housing 52 may be an annular housing which is housed between the internal wall 56 and the external wall 58 of the fairing 46 outside the part upstream 50 of fairing, that is to say preferably in a downstream part 66 situated downstream of said upstream part 50, and which is capable of being moved axially upstream to close off the torn part 54.
    However, in the preferred embodiment of the invention, as shown in FIG. 5, the palliative casing 52 comprises a plurality of angular segments 60 of palliative casing, angularly distributed around the axis XX of the fairing 46, each of which is movable between a retracted position, shown in Figures 4 and 6, in which it is retracted into a housing 62 of the fairing 46 formed outside the upstream part 50 of said fairing 46, and a deployed position, shown in Figures 7A and 7B, in which it extends to the right of the upstream fairing part 50, and more particularly to the right of the torn part 54, substantially in the extension of the fairing 46, to close off the torn part 54.
    The perspective view with cutaway of FIG. 5 represents the downstream part 66 of the fairing 46 receiving, in its housing 62, which is arranged between the inner 56 and outer 58 walls of the fairing 46, the angular segments 60 of the palliative housing 52.
    As can be seen, the angular segments 60 are distributed angularly in a regular manner around the axis XX of the fairing 46. They are here represented in the retracted position, it being understood that their deployed position corresponds to a displacement of these segments 60 towards the 'upstream, that is to the left of Figure 5.
    To ensure the passage of the angular sectors 60 from their retracted position to their deployed position, the invention proposes a deployment means 64 which has been shown diagrammatically in FIGS. 4, 6 7A and 7B. Any deployment means known from the state of the art can be used to move these angular segments 60 of palliative housing 52.
    For example, it is possible to associate one or more segments 60 with a hydraulic or electric actuator, or to associate them with a pyrotechnic means allowing rapid deployment thereof. It is also possible, as a variant, to subject these segments 60 to a restoring force provided by a spring (not shown) and to clamp these segments under normal flight conditions. In this case, the deployment of the segments 60 is ensured by releasing the clamping means of these segments 60 in order to allow the spring to be decompressed.
    Furthermore, it will be understood that the segments 60 are mounted on sliding guide means (not shown) which make it possible to guide them in sliding between the interior 56 and exterior 58 walls of the fairing 46 during their movements towards the upstream part 50. of the fairing 46. These sliding means can take several forms and, not being the subject of the present invention, they will not be described more explicitly in the present description.
    In the preferred embodiment of the invention, the aircraft comprises a palliative housing system comprising the palliative housing 52, as well as control means (not shown) of the palliative housing 52. These means are configured to be activated at least in response to blade break information 42 from the upstream fan 38, which is provided by an upstream blade break detection means (not shown) capable of providing said information. The means for controlling the palliative housing include the means for deploying the angular segments 64.
    In its simplest configuration, according to a first embodiment of the invention, when the upstream blade break detection means provides information representative of the break of a blade 42 of the upstream fan 38, the control means operate the deployment means of all sectors 60 simultaneously. This configuration has been shown in FIG. 7A. Only the sectors 60 corresponding to the torn part 54 effectively seal this torn part, the remaining sectors 60 being moved inside the intact part of the upstream part 50 of the fairing 46.
    In this case, as illustrated in FIG. 8, a method for controlling the palliative housing system comprises a first step ET1 during which the upstream blade break detection means 42 detects the break of at least one blade 42 of the upstream fan 38, and a second step ET2, determined by the first step ET1, during which the control means control the deployment of the palliative housing 52 and more particularly of all of these angular sectors 60.
    According to a second embodiment of the invention, the palliative case control system comprises means of identification (not shown) of the angular position of any torn part 54 of the upstream part 50 of the fairing 48. Thus, these means identification are likely to identify which angular zone of the upstream part 50 of the fairing 52 is actually deteriorated by a tear and comprises a torn part 54, and to deliver information representative of this position making it possible to carry out a selective control of the sectors angular 60. Thus, on the basis of this position information of any torn part 54, the control means of the palliative housing selectively activate the deployment means 64 of each angular sector 60 associated with a torn part 54 of the upstream part 50 of fairing 52, which avoids untimely deployment of sectors 60 for which a deployment is not necessary and therefore to save the energy corresponding to their deployment.
    This configuration proves to be particularly advantageous when the deployment of the angular sectors 60 is ensured by deployment means 64 of pyrotechnic origin which require replacement as soon as they have been used. Once the aircraft 10 has returned to its base, the turbojet engine 20 undergoes a maintenance operation aimed at replacing the torn part 54 of the upstream fairing part 50, the damaged elements of the upstream fan 38, and only the pyrotechnic deployment means 64 which have been used.
    In this case, as illustrated in FIG. 9, a method of controlling the palliative housing system comprises a first step ET1 during which the upstream blade break detection means 42 detects the break of at least one blade 42 of the upstream fan 38, a second step ET2 during which the means for identifying the angular position of any torn part 54 of the upstream fairing part 50 detect the position of any torn part 54 of the upstream part 50 fairing 46, and a third step ET3, determined by the first and second steps ET1 and ET2, during which the control means selectively activate the deployment means 64 of each angular sector 60 associated with a torn part 54 of the part upstream 50 of fairing.
    The invention is particularly applicable to a turbojet engine 20 of the type described above comprising two counter-rotating rotors 34, 36 respectively driving two coaxial and counter-rotating blowers upstream 38 and downstream 40 each comprising a crown of blades 42, 44, the all of the blowers 38, 40 is integrated into a fairing 46 of the turbojet engine 20. Although the configuration of the turbojet engine 20 is not limited to the invention, it will preferably be implemented in the context of an aircraft 10 including the turbojet engine 20 is arranged behind the fin 17 and behind the fuselage 12 so as to avoid any risk of interaction of a debris 43 with a vital part of the aircraft 10.
    It will also be noted that the palliative casing system may preferably be associated with control means intended to ensure the operability of the downstream fan 40, these control means aiming in particular to deactivate the upstream fan
    38 and to ensure satisfactory operation of the downstream blower 40. However, since these control means are not the subject of the invention, this is not discussed further in the following description.
    权利要求:
    Claims (11)
    [1" id="c-fr-0001]
    Claims
    1. Aircraft (10) comprising a fuselage (12), flight control surfaces (16, 18), and a turbojet engine (20) integrated at the rear of said fuselage (12) in extension thereof, the turbojet engine ( 12) comprising at least two gas generators (22) which supply, by a common central stream (30), a power turbine (32) comprising two counter-rotating rotors (34, 36) respectively driving two upstream coaxial and counter-rotating blowers (38 ) and downstream (40) each comprising a crown of blades (42, 44), all of the blowers (38, 40) being integrated in a fairing (46) of the turbojet engine (20) formed at the rear of the fuselage ( 12), characterized in that at least said fairing (46) is arranged axially behind the flight control surfaces (16, 18) and comprises an upstream part (50), surrounding the upstream blower (38), configured to be traversed radially by at least one fragment (43) of blade (42) of the upstream fan (38), in the event of r vane break (42) of said upstream fan (38) and ejection of said at least fragment (43).
    [2" id="c-fr-0002]
    2. Aircraft (10) according to the preceding claim, characterized in that the upstream part (50) of fairing (46) is configured to be at least partially torn by the passage of said at least blade fragment (43) during its crossing of said upstream fairing part (50) (46).
    [3" id="c-fr-0003]
    3. Aircraft (10) according to the preceding claim, characterized in that the fairing comprises at least one palliative housing (52) which is capable of being deployed on any torn part (54) of the upstream part (50) of fairing (46 ).
    [4" id="c-fr-0004]
    4. Aircraft (10) according to the preceding claim, characterized in that the palliative casing (52) comprises a plurality of angular segments (60) of the palliative casing, angularly distributed around the axis (XX) of the fairing (46), each of which is movable between a retracted position in which it is retracted in a housing (62) of the fairing (46) formed outside the upstream part (50) of said fairing, and a deployed position in which it extends in line with the upstream part (50) of fairing (46), substantially in the extension of the fairing (46), for closing said torn part (54).
    [5" id="c-fr-0005]
    5. Aircraft (10) according to one of claims 3 or 4, characterized in that it comprises a palliative housing system comprising the palliative housing (52) and means for controlling the palliative housing (52) which are configured for be activated at least in response to information on the failure of a blade (42) from the upstream fan (38), and means for detecting failure of a blade upstream (42) capable of providing said information.
    [6" id="c-fr-0006]
    6. Aircraft (10) according to claims 4 and 5 taken in combination, characterized in that the control means of the palliative housing (52) comprise a deployment means (64) of each angular segment (60).
    [7" id="c-fr-0007]
    7. Aircraft (10) according to the preceding claim, characterized in that the palliative case control system (52) comprises means for identifying the angular position of any torn part (54) of the upstream part (50) of fairing (46) capable of providing position information of any torn part (54), and in that the control means of the palliative housing (52) are configured to selectively activate the deployment means (64) of each angular sector ( 60) associated with a torn part (54) of the upstream fairing part (50) (46) in response to said position information.
    [8" id="c-fr-0008]
    8. Turbojet engine (20) comprising two counter-rotating rotors (34, 36) respectively driving two upstream (38) and downstream (40) coaxial and counter-rotating blowers each comprising a crown of blades (42), all of the blowers (38, 40) being integrated in a fairing (46) of the turbojet engine (20), characterized in that said fairing (46) comprises at least one palliative casing (52) which is capable of being deployed on any part (50) of the upstream fairing susceptible to be traversed radially and at least partially torn by the rupture of a blade (42) of the upstream fan (38), to close off the corresponding torn part (54).
    [9" id="c-fr-0009]
    9. A method of controlling an aircraft pallet housing system according to one of claims 5 to 7, characterized in that it comprises a first step (ET1) during which the means for detecting rupture of upstream blade detects the rupture of at least one blade (42) of the upstream blower (38), and a second step (ET2), determined by the first step (ET1), during which the control means control the deployment of the palliative housing (52).
    [10" id="c-fr-0010]
    10. Method for controlling an aircraft pallet housing system
    5 according to claim 7, characterized in that it comprises a first step (ET1) during which the upstream blade break detection means detects the break of at least one blade (42) of the upstream fan ( 38), a second step (ET2) during which the means for identifying the angular position of any torn part (54) of the upstream part (50) of
    10 fairing (46) detect the position of any torn part (54) of the upstream part (50) of fairing (46), and a third step (ET3), determined by the first and second steps (ET1, ET2), at during which the control means selectively activate the deployment means (64) of each angular sector (60) associated with a torn part (54) of the upstream part (50)
    [11" id="c-fr-0011]
    15 of fairing (46).
    1/5 x
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    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    EP0184962A1|1984-12-06|1986-06-18|Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A."|Retention housing for a turbo fan|
    FR2997681A1|2012-11-08|2014-05-09|Snecma|PLANE PROPELLED BY A TURBOREACTOR WITH CONTRAROTATIVE BLOWERS|
    FR3030445A1|2014-12-22|2016-06-24|Airbus Operations Sas|AIRCRAFT TURBOMACHINE PROPELLER, COMPRISING A BLADE RETENTION STRUCTURE CONNECTED TO THE EXTERNAL RADIAL END OF EACH BLADE|WO2019243117A1|2018-06-21|2019-12-26|Safran|Rear propulsion system for an aircraft|CN103429495B|2011-02-11|2017-04-12|空中客车运营简化股份公司|Airplane having a rear propulsion system|FR3060531B1|2016-12-20|2019-05-31|Airbus Operations|REAR AIRCRAFT PART COMPRISING A FUSELAGE FRAME SUPPORTING TWO PARTIALLY BITTED ENGINES|
    US10723470B2|2017-06-12|2020-07-28|Raytheon Technologies Corporation|Aft fan counter-rotating turbine engine|
    US10737797B2|2017-07-21|2020-08-11|General Electric Company|Vertical takeoff and landing aircraft|
    DE102018208297A1|2018-05-25|2019-11-28|Rolls-Royce Deutschland Ltd & Co Kg|Aircraft with at least one jet engine|
    法律状态:
    2017-05-02| PLFP| Fee payment|Year of fee payment: 2 |
    2018-02-02| PLSC| Search report ready|Effective date: 20180202 |
    2018-06-21| PLFP| Fee payment|Year of fee payment: 3 |
    2019-06-21| PLFP| Fee payment|Year of fee payment: 4 |
    2020-06-23| PLFP| Fee payment|Year of fee payment: 5 |
    2021-06-23| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1657186|2016-07-26|
    FR1657186A|FR3054526B1|2016-07-26|2016-07-26|AIRCRAFT COMPRISING AN INTEGRATED REAR-FUSELAGE TURBOREACTOR COMPRISING A FITTING FOR EJECTING OF BLADES|FR1657186A| FR3054526B1|2016-07-26|2016-07-26|AIRCRAFT COMPRISING AN INTEGRATED REAR-FUSELAGE TURBOREACTOR COMPRISING A FITTING FOR EJECTING OF BLADES|
    US15/659,546| US10746044B2|2016-07-26|2017-07-25|Aircraft comprising a turbojet engine integrated into the rear fuselage comprising a fairing allowing the ejection of blades|
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