• 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 proposes a new type of turbojet engine to overcome the problem of UERF or any other equivalent problem without resorting to the installation of a shield. The engine (1, 1 ') turbojet engine comprises at least three zones including an air intake zone and an exhaust zone. The axis of the air intake zone is not coincident with the axis of the exhaust zone of said engine (1, 1 '), the engine (1, 1') thus having at least two axes intersecting and being called multi-axial motor. Thus, the turbojet engine has at least two zones of different longitudinal orientations: by choosing the orientation of engine zones more sensitive to the detachment or breaking of elements the gas generator for example, it also allows to choose the direction possible trajectories of these detached elements.
    公开号:FR3045011A1
    申请号:FR1662397
    申请日:2016-12-14
    公开日:2017-06-16
    发明作者:Julien Guillemaut；Olivier Cazals
    申请人:Airbus Operations SAS；
    IPC主号:
    专利说明:

    The present invention relates to the field of turbojets and their arrangement in the rear part of an aircraft. The present invention relates to a new type of turbojet, the rear part of the aircraft carrying the said turbojet engines, the method of producing said rear part and the aircraft having such a rear part. It applies more particularly to commercial aircraft.
    TECHNICAL AREA
    The present invention is concerned with aircraft equipped with two turbojet engines still called turbofan installed in the rear part of the fuselage on either side of it.
    The rear part of the fuselage comprises a portion of variable section carrying the empennage and located at the rear of the aircraft is opposite the cockpit in a conventional type configuration.
    PRIOR ART
    The application EP 11382409.8 filed on December 28, 2011 by the present applicant describes an aircraft in which two engines are arranged at the rear of both sides of the fuselage. FIG. 3 illustrates, for example, turbojet engines half-buried on either side of a plane of symmetry of the fuselage. Engines of the semi-buried type have the advantage of being able to ingest part of the boundary layer and to improve its performance.
    However, because of being partially buried, the distance between them decreases. The risk of being impacted by an Uncontained Engine Rotor Failure (UERF) event increases. The event type UERF is characterized by the detachment of an internal part of the turbojet engine that hits the fuselage or the opposite engine directly or indirectly. One solution to avoid it proposed by the application in question consists in providing an internal shield positioned in a plane of vertical symmetry of the fuselage.
    However the addition of a shield with a sufficiently solid structure increases the weight of the aircraft. It also requires in view of its size to review the internal organization of the tail of the aircraft to allow its installation.
    The applications WO2014 / 074149 and US2014 / 025216 describe an air inlet zone, an exhaust zone, the axis of the air inlet zone not being merged with the axis of the zone d 'exhaust. In the application WO2014 / 074149, the air inlet zone corresponds to a simple opening in the fuselage of the aircraft. In the application US2014 / 0252161, the two parts of the axis motor not confused are connected by an axis which drives propellers located in the axis of the exhaust zone.
    The present invention aims to propose a new type of turbojet engine offering an alternative to increase the ingestion of the boundary layer and therefore the engine performance while overcoming the problem of the FRAU or any equivalent problem without resorting to the installation of a shield.
    SUMMARY OF THE INVENTION
    To do this, the present invention proposes a turbojet engine comprising at least three zones including an air intake zone and an exhaust zone characterized in that the axis of the air intake zone is not coincide with the axis of the exhaust zone of said engine, the motor thereby having at least two axes intersecting and being called multi-axial motor.
    Thus, the turbojet engine has at least two parts of different longitudinal orientations: by choosing the orientation of engine areas more sensitive to the detachment or breaking of elements the gas generator for example, it also allows to choose the direction possible trajectories of these detached elements. Thus when the engine will be placed on the rear part of an aircraft at its variable section, having chosen an appropriate orientation for certain areas of the engine, the direction of the trajectories of the detached elements does not meet the arranged motor of the other side of the rear part of the fuselage.
    The turbojet engine has at least one of the following optional features, taken alone or in combination. The axis of the air inlet zone and the axis of the exhaust zone are parallel.
    The turbojet engine comprises a training zone; the axis of the drive zone is neither parallel nor coincident with the axes of the air inlet zone and the exhaust zone, the engine thus having three different axes.
    The turbojet engine comprises a compression zone and a combustion zone; the axis of the training zone, compression and combustion are combined.
    The present invention also relates to an aircraft rear portion having a fuselage portion of variable section comprising at least two turbojet engines positioned on either side of said portion and comprising at least three zones including an air intake zone. and an exhaust zone, the axis of the air inlet zone not being merged with the axis of the exhaust zone of said engine, the engine thus having at least two intersecting axes and being called multi-axial motor characterized in that the axis (s) of other areas of the multi-axial turbojet engine is (are) oriented so that one or more surface (s) of delimitation trajectories detached elements of each turbojet engine does not meet the opposite turbojet engine.
    The aircraft aft portion has at least one of the following optional features, taken alone or in combination.
    Said surface consists of a cone representative of a UERF event established for a driving zone of said engine.
    The shape of the fuselage and / or the shape and positioning of the various means for fixing the turbojet engines to the fuselage are determined to allow the zone or areas included between the air inlet zone and the air intake zone. exhausting said engines to follow the contour of the fuselage and to orient the delimiting surface or surfaces.
    The exhaust zone of the engine and the exhaust zone of the other engine have merged to form a single exhaust zone positioned at the rear end of said portion.
    The exhaust zone is provided with a thrust reversal system.
    The present invention also relates to the aircraft provided with such a rear part.
    The present invention also relates to a method for producing a rear part of an aircraft having a variable section, carrying at least two turbojet engines comprising at least three zones including an air intake zone and an exhaust zone, the axis of the air intake zone not being confused with the axis of the exhaust zone of said engine, the engine having at least two intersecting axes and being called a multiaxial motor, the method being characterized in that it comprises a step of positioning the motors on either side of said variable portion so that one or more surface (s) for delimiting the trajectories of detached elements of each turbojet engine do not encounter the opposite turbojet engine.
    The method comprises a step of choosing the orientation of the axes of the multi-axial motors and to play on the shape of the fuselage and / or the shape and positioning of the various means of fixing the turbojet engines to the fuselage to help guide the or bounding surfaces and motors relative to the fuselage contour.
    SUMMARY DESCRIPTION OF THE DRAWINGS Other objects, advantages and characteristics of the invention will appear on reading the following description of the turbojet engine and the rear part of an aircraft provided with such an engine according to the invention, given by way of non-limiting example with reference to the accompanying drawings in which: • Figure 1 shows a schematic side sectional view of a known type of turbojet engine; • Figure 2 shows a schematic top sectional view of an aircraft rear portion provided on either side of two turbojet engines according to one embodiment of the invention; • Figure 3 shows a schematic top sectional view of an aircraft rear portion provided on either side of two turbojet engines according to another embodiment of the invention; • Figure 4 shows a schematic view in section from above of an aircraft rear portion provided on either side of two turbojet engines according to another embodiment of the invention; FIGS. 5a to 5d show a comparison of the impact of a UERF event on a rear part according to the prior art and on a rear part according to the embodiments of FIGS. 2 and 3.
    MANNER OF REALIZING THE INVENTION
    As shown in FIG. 1, the present invention relates to an engine 1 turbojet engine or turboprop engine in which air is sucked and compressed to be then mixed with a fuel whose combustion causes a large expansion of the gases: exhaust gas provides the thrust allowing the movement of the aircraft forward but also allowing the movement of the compressor performing said compression.
    Throughout the following description, by convention, the X-X direction corresponds to the longitudinal direction of the aircraft, which is comparable to the longitudinal direction of the rear portion 2 thereof. On the other hand, the terms "front" and "rear" are to be considered in relation to a direction of advancement of the aircraft encountered following the thrust exerted by the turbojet engine (s) 1, this direction being represented schematically by the arrow 4.
    The turbojet engine 1 comprises at least five zones: an air intake zone 6 comprising an air inlet 8 which directs the penetration of air into the engine represented by the arrows 10 and in which is housed a propeller 12 called blower for the suction of air; - A compression zone 14 provided with a compressor 16 for gradually increasing the pressure of the sucked air; a zone 18 of combustion including a combustion chamber 20 in which the fuel is injected into the compressed air causing its combustion and the violent ejection rearward of hot gases represented by the arrows 22; a drive zone 24 comprising a turbine 26 driven by the ejection 22 of hot gases and in turn allowing the propeller 12 and the compressor 16 to move to which it is linked by an axis 28; an exhaust zone 30 having an exhaust nozzle 32 regulating the exit of the gases 22 providing the thrust allowing the movement of the aircraft forward represented by the arrow 4.
    In the following description, will be considered as axis of a zone, the central longitudinal axis of partial or total symmetry of the components or a part of said component of this area. In the case where a zone has no component having a central axis of partial or total symmetry, consider the axis of a neighboring zone.
    Thus, for example, the axis A-A of the air intake zone 6 in the examples illustrated in FIGS. 1 to 5 is not an axis of symmetry of the nacelle 34 at the level of the air inlet. The engine has a semi-buried configuration. In such a configuration, part of the nacelle 34 of the turbojet engine 1 is constituted by the fuselage 36 and thus the nacelle 34 does not have as in the engine of Figure 1 a symmetrical shape. As a result, the axis AA of the air intake is constituted by the axis of rotation of the fan 12 since it constitutes an axis of symmetry for said fan 12. The axis BB of the compression zone 14 is constituted by the axis of the compressor 16 and more precisely the axis of rotation of the blades 38 (blades or equivalents) that it carries. The C-C axis of the combustion zone 18 is the longitudinal axis of symmetry of the combustion chamber. In the case where the chamber has a shape devoid of central longitudinal axis of symmetry, the axis of the combustion chamber is the axis of the compression zone 14 and / or the zone 24 of training. The axis DD of the zone 24 for driving is constituted by the axis of the turbine 26 and more precisely by the axis of the blades 40 of said turbine 26. The axis EE of the zone 30 of exhaust is constituted by the axis of the outlet of the nozzle 32.
    In the known type of engines, the axes AA of the air inlet zone, BB of the compression zone, CC of the combustion zone, DD of the entrainment zone and EE of the exhaust zone are all confounded along one and the same axis FF as shown in Figure 1. The zones follow one another and are centered around the same longitudinal axis.
    To provide a new engine configuration to address the problems discussed above, the axis of the air intake zone 6 is not confused with the axis of the exhaust zone. Therefore, if the axes of these extreme zones are not confused, it follows that there is at least one zone whose axis intersects at least one of the axes of the extreme zones to allow the connection. As a result, the turbojet engine is said to be multiaxial because the different zones that constitute it have at least two different non-coinciding axes that intersect. The turbojet engine 1 does not have a slender shape centered on a single axis (F-F in the illustrated prior art). The different components of the turbojet engine are not centered on one and the same axis. The axis of one or more zones is different from the axis of one or more other zones. The motor has at least two zones oriented in a different longitudinal direction.
    The engine has areas at which elements can break or detach account given for example vibration or other thermomechanical effects produced during operation of the engine at these areas. Thus, for example, the rotational movement of the turbine of a very high speed motor can cause break-off or take-off or other detachment of elements, pieces, debris or equivalent, hereinafter referred to as loose elements. The analysis of these zones leads to the identification of the trajectories followed by these detached elements. Thus for the turbine of a turbojet engine, for example, it is known that the detached elements are contained in a geometric surface of conical shape called cone. Said surface could have any other type of shape and will be generally referred to in the following surface delimiting trajectories of detached elements.
    According to an embodiment such as those illustrated in FIGS. 2 to 5, the axes A-A of the zone of the air inlet and E-E of the exhaust zone are parallel but not coincidental. The axis of the driving zone 24 is neither parallel nor coincident with the axes of the air inlet zone 6 and the exhaust zone. Only the B-B axes of the compression zone 14, C-C of the combustion zone 18 and D-D of the entrainment zone coincide. The axes A-A and E-E on the one hand and B-B, C-C and D-D on the other hand intersect and form an angle different from 90 ° or 180 °.
    The present invention relates to the field of aircraft whose rear portion 2 has a variable section. The rear part 2 of the aircraft according to the invention and shown diagrammatically in FIGS. 2 to 5 has a central longitudinal axis X-X through which passes a vertical plane of symmetry when the aircraft is on the ground in a horizontal position. The rear part 2 carries two turbojets 1,1 'disposed on either side of the plane of symmetry passing through the longitudinal axis X-X.
    In all the configurations illustrated in FIGS. 2 to 5, the turbojet engines 1, 1 'are positioned on either side of the rear part 2 of the aircraft along axes AA of the zone of the inlet of FIG. air and EE of the exhaust zone parallel to the longitudinal axis XX of the rear portion 2 of the aircraft. As a result, the air sucked by said turbojet engines 1, 1 'is rejected parallel to the air sucked in, and the air sucked in and ejected by the other turbojet engine respectively 1', 1 and parallel to the axis of the aircraft ensuring a movement in rectilinear translation according to the arrow 4.
    As shown in Figures 2 to 5, the turbojet engines 1, 1 'are positioned along the rear portion of the aircraft of variable section. In order to be able to follow the lines of the fuselage 36 and determine the orientation of the boundary surface (s) of the detached element trajectories, the turbojet engine 1, 1 'is a multiaxial motor as presented previously.
    One or more zones of the turbojet engine corresponding to the at least one sensitive zone of said engine are positioned along one or more axes making it possible to orient the delimiting surface or surfaces so that they do not meet the opposite turbojet engine.
    It is also possible to play on other parameters such as the shape of the fuselage and more precisely the curvature of the variable section or the shape, and in particular the length, of the different means of attachment of the turbojet to the fuselage or their positioning on this one.
    According to the embodiments shown in FIGS. 2 to 5, the axis of the air inlet zone 6 is parallel but not coincidental with that of the exhaust zone and these two axes AA and EE are neither parallel or confused with those of other areas. Each axis B-B, C-C and D-D respectively intersects the axes A-A and E-E.
    In this way, the air inlet zone 6 is substantially parallel to the X-X axis of the rear portion 2 and may be positioned closest to it to increase the ingestion of the boundary layer.
    As shown in FIGS. 5a and 5b, each turbojet engine 1, 1 'has a bounding surface of the detached component trajectories representative of a UERF event in the form of a cone 41. The cone 41 defines, as previously seen, the surface inside which are all the various possible trajectories followed by detached elements of the turbojet and in particular of the turbine 26. In the prior art, as shown in FIGS. 5a and 5b, the positioning of the cone 41 is such that parts of a turbojet 1 could come to hit the turbojet 1 'opposite.
    In the present invention, as shown in FIGS. 5c and 5d, the axis DD of the drive zone 24 is oriented in such a way that the cone 41 representative of a UERF event for each of the turbojet engines 1, 1 'does not cut not, do not cross the other turbojet engine 1,1 '. Thus, in case of detachment or even rupture of an element of the turbine and / or an element of the blades of said turbine of a turbojet respectively 1.1 ', the detached elements can not damage or destroy the other turbojet engine 1 ', 1.
    The shape of the cone 41 representative of a UERF event depends on the turbojet engine 1. Depending on the shape of the cone, the axis of the driving zone 24 of the corresponding turbojet engine 1 or 1 'is determined in such a way that said cone does not meet the other turbojet respectively 1 'or 1 and is therefore positioned totally in front of said turbojet engine. It is also possible to play as seen previously on other parameters such as the shape of the rear part as the curvature of the variable section 42 of the fuselage or the shape, and in particular the length, of the different means 43, 43 ', 44, 45, 44 ', 45' of fixing the turbojet to the fuselage or their positioning thereon. All of these parameters are chosen to allow positioning the cone as desired while positioning the turbojet engine along the rear portion of variable section.
    In all of the embodiments illustrated in FIGS. 2 to 5, the attachment means of the turbojet engine 1 and 1 'on the fuselage are in the form of three fasteners respectively (43, 44, 45) and (43', 44 ', 45').
    The first attachment allows via a connection 43, 43 'the direct attachment of the zone 24 for driving the turbojet engine 1, 1' to the rear portion 42 of variable section of the fuselage. The fasteners 43, 43 'are connected by a rod 46 passing through the internal rear portion 42 of variable section of the fuselage.
    The second attachment allows the attachment via a connecting rod 44, 44 'of the compression zone 14 of the turbojet engine 1, 1' to the rear portion 42 of variable section of the fuselage.
    The third fastener allows the attachment by means of a connecting rod 45, 45 'of the air intake zone 6 of the turbojet engine 1, 1' to the rear portion 42 of variable section of the fuselage.
    In the embodiment of Figure 2, the attachment of the first attachment is at a frame of the fuselage. The attachment of the third attachment is also at a frame of the fuselage.
    The second attachment can be removed: it makes it possible to reinforce the maintenance of the turbojet engine.
    The additional and distinctive features of the embodiments of FIGS. 3 and 4 with respect to that of FIG. 2 are as follows: the exhaust zones of the engines 1, 1 'turbojet engine have merged to become one: the zone exhaust is therefore positioned at the rear end of the rear portion on the axis XX thereof. The E-E axis of the exhaust zone 30 coincides with the X-X axis of the rear part.
    This allows to have only one nozzle instead of two. It follows a gain in weight, size, cost of manufacture, maintenance ...
    In the embodiment of FIG. 4 with respect to that of FIG. 3, the zone 30 is provided with a thrust reverser system 47 which again makes it possible to obtain the advantages enumerated above. The thrust reversal system is a system of known type, for example in the form of two flaps articulated on the edge of the nozzle of the exhaust zone.
    权利要求:
    Claims (11)
    [1" id="c-fr-0001]
    1 - turbojet engine comprising at least three zones including an air intake zone (6) and an exhaust zone (30) characterized in that the axis of the air intake zone (6) is not confused with the axis of the exhaust zone (30) of said engine (1, T), the engine (1, T) thus having at least two intersecting axes and being called a multiaxial motor, in that the motor comprises a zone (24) for driving and in that the axis of the zone (24) for driving is neither parallel nor coincident with the axes of the input zone (6). air and the zone (30) exhaust, the engine (1, T) thus having three different axes.
    [2" id="c-fr-0002]
    2 - turbojet engine according to claim 1, characterized in that the axis of the zone (6) of air inlet and the axis of the zone (30) exhaust are parallel.
    [3" id="c-fr-0003]
    3 - turbojet engine according to one of claims 1 or 2, characterized in that it comprises a zone (14) of compression and a zone (18) of combustion and in that the axis of the zone (24) drive, (14) compression and (18) combustion are combined.
    [4" id="c-fr-0004]
    4 - Rear part of aircraft having a portion (42) fuselage of variable section comprising at least two engines (1, T) turbojet engines according to one of claims 1 to 3 positioned on either side of said portion (42). ), characterized in that the axis (s) of other zones of the multi-axial turbojet engine is (are) oriented so that one or more surface (s) of delimitation of the trajectories of detached elements of each turbojet engine (1, T) does not encounter the opposite turbojet engine (T, 1)
    [5" id="c-fr-0005]
    5 - Rear part of aircraft according to claim 4, characterized in that the surface consists of a cone (41) representative of a UERF event established for a zone (24) for driving said motor.
    [6" id="c-fr-0006]
    Aircraft rear part according to one of Claims 4 or 5, characterized in that the shape of the fuselage and / or the shape and positioning of the various means (43, 43 ', 44, 45, 44', 45 ') fixing the turbojet engines (1, 1') to the fuselage are determined to allow the zone or zones between the air inlet zone and the exhaust zone of said engines to follow the contour fuselage and orient the boundary surface (s).
    [7" id="c-fr-0007]
    7 - Rear part of an aircraft according to one of claims 4 to 6, characterized in that the zones (30) of the engine exhaust (1) and (1 ') have merged to form a single zone d exhaust positioned at the rear end of said portion
    [8" id="c-fr-0008]
    8 - Rear part of aircraft according to claim 7, characterized in that the exhaust zone is provided with a thrust reversal system.
    [9" id="c-fr-0009]
    9 - Aircraft characterized in that it comprises a rear portion according to one of claims 4 to 8.
    [10" id="c-fr-0010]
    10 - Method of producing a rear part of an aircraft carrying at least two engines (1, 1 ') turbojet engines having a variable section (42) and comprising at least three zones including an input zone (6) of air and an exhaust zone (30), the axis of the air intake zone (6) not being merged with the axis of the exhaust zone (30) of said engine (1, 1 '), the motor (1, 1') thereby having at least two intersecting axes and being referred to as a multiaxial motor characterized in that the motors comprise a driving zone (24) and in that the axis of the driving zone (24) is neither parallel nor confused with the axes of the zone (6) of air intake and the zone (30) of exhaust, the motors (1, 1 ') thus having three different axes and in that the method consists in positioning the motors (1, 1 ') on either side of said variable portion (42) so that one or more surfac e (s) of delimitation of the trajectories of detached elements of each turbojet engine (1, 1 ') does not meet (not) the engine turbojet opposite (1', 1),
    [11" id="c-fr-0011]
    11 - Process according to claim 10, characterized in that it consists in choosing the orientation of the axes of the motors (1, 1 ') multi-axial and play on the shape of the fuselage and / or the shape and positioning of various means (43, 43 ', 44, 45, 44', 45 ') for attaching the turbojet engines (1, 1') to the fuselage to enable the delimiting surface (s) and the engines to be oriented relative to the outline of the fuselage.
    类似技术:
    公开号 | 公开日 | 专利标题
    FR3045011A1|2017-06-16|MULTI-AXIAL TURBOREACTOR AND REAR AIRCRAFT PART PROVIDED WITH SUCH TURBOJET ENGINES
    CA2890274C|2019-10-22|Aircraft propelled by a turbojet engine with contrarotating fans
    EP2488739B1|2018-09-12|Air intake for a as turbine engine within a nacelle
    CA2732664C|2017-01-03|Airplane having engines partially encased in the fuselage
    FR2873096A1|2006-01-20|MULTI-ENGINE
    EP2414655B1|2013-06-05|Rotating inlet cone for a turbomachine including an eccentric front end and corresponding turbomachine
    FR2958624A1|2011-10-14|AIRCRAFT TURBOHOMOTING MACHINE, COMPRISING AN INCOMING MOBILE REAR FLAP
    CA2694996C|2015-10-06|Noise control chevron for a nozzle, and nozzle and turboshaft engine provided with such a chevron
    FR2935349A1|2010-03-05|TURBOMACHINE WITH NON-CARINEATED PROPELLERS
    FR2926789A1|2009-07-31|NACELLE FOR TURBOREACTOR
    CA2798679C|2018-03-27|Device for reducing the noise emitted by the jet of an aircraft propulsion engine
    FR2983834A1|2013-06-14|HOOK PYLONE FOR TURBOMACHINE
    FR2610592A1|1988-08-12|COMBINATION OF AIRPLANE AND DRIVE GROUP
    CA2576700C|2014-07-08|Bypass engine air inlet
    WO2017060585A1|2017-04-13|Aircraft with multiple fan propulsion assembly fixed under the wing
    CA2716742A1|2009-09-17|Attachment structure for a turbojet engine
    EP3380399B1|2021-04-28|Aircraft powered by a turbomachine provided with an acoustic baffle
    FR3033545A1|2016-09-16|NACELLE FOR AIRCRAFT ENGINE ASSEMBLY COMPRISING AT LEAST ONE ARTICULATED NACELLE COVER AT ITS FRONT END
    EP3325771A1|2018-05-30|Aircraft comprising two contra-rotating fans to the rear of the fuselage, with spacing of the blades of the downstream fan
    EP3540205A1|2019-09-18|Aircraft drive unit in which the nacelle is linked by a pivot to the drive shaft of its fan
    FR2951504A1|2011-04-22|Gas turbine engine and nacelle assembly for e.g. helicopter, has secondary deflecting channel shaped such that flow velocity of air increases from upstream to downstream, where channel has outlet with opening leading into wall of nacelle
    WO2010061071A2|2010-06-03|Nacelle integrated on a flying wing
    FR3020338A1|2015-10-30|REAR AIRCRAFT PART PROVIDED WITH SUPPORT STRUCTURE OF OPTIMIZED SHAPE ENGINES
    FR2993861A1|2014-01-31|Multi-flow turbojet engine assembly for aircraft, has nacelle comprising upstream air inlet sleeve with air intake in region of leading edge, where shield of pylon is extended up to air intake of air inlet sleeve
    FR2986504A1|2013-08-09|Device for attaching turbopropeller with structure element of aircraft, has center line intersected with driving axis in point along radial direction, where portion of center line forms angle with radial direction
    同族专利:
    公开号 | 公开日
    US20170167437A1|2017-06-15|
    FR3045010A1|2017-06-16|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    FR2965250A1|2010-09-28|2012-03-30|Snecma|Installation for engines i.e. turbojet engines, at back of fuselage of twin-jet engine aircraft, has engines whose axles are directed in oblique manner with respect to longitudinal plane of aircraft for forming equal angles|
    WO2014074149A1|2012-11-12|2014-05-15|United Technologies Corporation|Stabilizer sacrificial surfaces|
    US20140252161A1|2013-03-11|2014-09-11|United Technologies Corporation|De-couple geared turbo-fan engine and aircraft|
    EP3048048A1|2015-01-21|2016-07-27|Rolls-Royce plc|An aircraft|
    FR1238410A|1959-07-02|1960-08-12|Method and devices for ensuring the stability of multi-jet airplanes in the event of failure of part of the engines|
    US3099425A|1960-12-16|1963-07-30|Hamburger Flugzeugbau Gmbh|Jet propulsion system|
    US3194516A|1962-10-22|1965-07-13|Messerschmitt Ag|Arrangement for jet engines at the tail end of aircraft|
    US3251567A|1963-03-25|1966-05-17|Messerschmitt Ag|Mounting of dual cycle propulsion units in the tail of an aircraft|
    FR1472962A|1964-08-08|1967-03-17|Dornier Werke Gmbh|engine equipment for airplanes|
    US6575406B2|2001-01-19|2003-06-10|The Boeing Company|Integrated and/or modular high-speed aircraft|
    GB2400411B|2003-04-10|2006-09-06|Rolls Royce Plc|Turbofan arrangement|
    US9309781B2|2011-01-31|2016-04-12|General Electric Company|Heated booster splitter plenum|
    US8789354B2|2012-02-10|2014-07-29|United Technologies Corporation|Gas turbine engine with separate core and propulsion unit|
    WO2014074135A1|2012-11-12|2014-05-15|United Technologies Corporation|Reverse core turbine engine mounted above aircraft wing|
    WO2014092757A1|2012-12-11|2014-06-19|United Technologies Corporation|Asymmetric thrust reversers|
    US9726112B2|2013-03-07|2017-08-08|United Technologies Corporation|Reverse flow gas turbine engine airflow bypass|
    US20140252160A1|2013-03-07|2014-09-11|United Technologies Corporation|Reverse flow gas turbine engine removable core|
    US9644537B2|2013-03-14|2017-05-09|United Technologies Corporation|Free stream intake with particle separator for reverse core engine|
    US20150330300A1|2013-03-14|2015-11-19|United Technologies Corporation|Two spool engine core with a starter|
    WO2015006009A1|2013-07-08|2015-01-15|United Technologies Corporation|Angled core engine|
    US10156206B2|2013-10-24|2018-12-18|United Technologies Corporation|Pivoting blocker door|
    US10094281B2|2014-01-30|2018-10-09|United Technologies Corporation|Gas turbine engine with twin offset gas generators|
    US10024235B2|2014-03-03|2018-07-17|United Technologies Corporation|Offset core engine architecture|
    US9989011B2|2014-04-15|2018-06-05|United Technologies Corporation|Reverse flow single spool core gas turbine engine|
    US10421554B2|2015-10-05|2019-09-24|United Technologies Corporation|Double propulsor imbedded in aircraft tail with single core engine|
    US10267263B2|2016-08-08|2019-04-23|United Technologies Corporation|Exhaust duct for turbine forward of fan|
    US10428740B2|2016-12-08|2019-10-01|United Technologies Corporation|Twin shafts driving adjacent fans for aircraft propulsion|
    US20180163664A1|2016-12-08|2018-06-14|United Technologies Corporation|Concentric shafts driving adjacent fans for aircraft propulsion|FR3052743B1|2016-06-20|2018-07-06|Airbus Operations|AIRCRAFT ASSEMBLY COMPRISING PROPULSION ENGINES BY INGESTION OF THE LIMIT LAYER|
    FR3060531B1|2016-12-20|2019-05-31|Airbus Operations|REAR AIRCRAFT PART COMPRISING A FUSELAGE FRAME SUPPORTING TWO PARTIALLY BITTED ENGINES|
    US11111029B2|2017-07-28|2021-09-07|The Boeing Company|System and method for operating a boundary layer ingestion fan|
    US10759545B2|2018-06-19|2020-09-01|Raytheon Technologies Corporation|Hybrid electric aircraft system with distributed propulsion|
    US10906657B2|2018-06-19|2021-02-02|Raytheon Technologies Corporation|Aircraft system with distributed propulsion|
    法律状态:
    2017-08-11| PLSC| Publication of the preliminary search report|Effective date: 20170811 |
    2017-12-21| PLFP| Fee payment|Year of fee payment: 2 |
    2019-12-19| PLFP| Fee payment|Year of fee payment: 4 |
    2020-12-23| PLFP| Fee payment|Year of fee payment: 5 |
    2021-12-24| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1562341A|FR3045010A1|2015-12-15|2015-12-15|MULTI-AXIAL TURBOREACTOR AND REAR AIRCRAFT PART PROVIDED WITH SUCH TURBOJET ENGINES|
    [返回顶部]