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1.发明申请公开(公告)号：US20180023406A1
公开(公告)日：2018-01-25
申请号：US15653976
申请日：2017-07-19
发明人： Cédric Zaccardi , Christophe Paul Jacquemard , Thierry Georges Paul Papin , Christophe Marcel Lucien Perdrigeon
CPC分类号： F01D9/065 , F01D5/185 , F01D9/041 , F01D9/042 , F01D25/12 , F01D25/18 , F02C7/16 , F02C7/18 , F02K3/06 , F05D2220/32 , F05D2230/25 , F05D2230/53 , F05D2240/12 , F05D2260/205 , F05D2260/208 , F05D2260/213 , F05D2260/98 , F28D2021/0089 , Y02T50/676
摘要： The invention relates to an intermediate case (25) for a twin spool turbomachine for an aircraft, comprising a hub (26), an outer shell (23) and outlet guide vanes (24) installed at their ends on the hub and on the outer shell, and each of at least some of the outlet guide vanes (24) performing a heat exchanger function and comprising a lubricant passage (50a, 50b) designed to be cooled by the fan flow (58) following an outer surface of the outlet guide vane. According to the invention, the case also comprises at least one lubricant duct (55) passing along a circumferential direction of the hub (26) and at least part of which is made from a single casting with the hub, the lubricant duct (55) having at least one lateral opening communicating with the lubricant passage (50a, 50b) of at least one of the vanes (24).
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公开(公告)号：US10723476B2
公开(公告)日：2020-07-28
申请号：US16185257
申请日：2018-11-09
发明人： Cedric Zaccardi , Christophe Marcel Lucien Perdrigeon , Francois Marie Paul Marlin , Thierry Georges Paul Papin , Jacky Novi Mardjono
摘要： A ring of vanes for an aircraft turbine engine, the ring presenting an axial direction and a radial direction and including a first annular sheath presenting an inside surface and a second annular sheath presenting an inside surface facing the first inside surface of the first annular sheath, the first and second sheaths being coaxial and defining between them a flow passage for a gas stream. The ring further includes both partially annular acoustic treatment panels including resonant cavities and also vanes extending in the radial direction between the first and second sheaths. Each vane includes a radial airfoil with at least two attachment tabs at each radial end of the airfoil fastened to the first or second sheath, and the inside surface of at least one of the first and second annular sheaths and the corresponding attachment tabs are covered by acoustic treatment panels arranged between the airfoils.
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公开(公告)号：US10697312B2
公开(公告)日：2020-06-30
申请号：US15914150
申请日：2018-03-07
发明人： Mohamed-Lamine Boutaleb , Fabien Roger Gaston Caty , Sebastien Vincent Francois Dreano , Thierry Georges Paul Papin , Christophe Marcel Lucien Perdrigeon , Cedric Zaccardi
IPC分类号： F01D9/06 , F01D9/04 , F01D25/12 , F01D25/18 , F02C7/14 , F28D21/00 , F01D5/18 , F28F1/40 , F28D1/02 , F28D1/053
摘要： A guide vane for a twin-spool aircraft turbomachine has an aerodynamic part that includes an internal lubricant cooling passage extending along a principal lubricant flow direction. The aerodynamic part is made in a single piece and also includes heat transfer fins arranged in the passage connecting the intrados and extrados walls and extending approximately parallel to the direction, these fins being distributed in successive rows along the principal direction and made such that for two rows of staggered directly consecutive fins, a first row includes fins forming a positive acute angle A1 with a dummy reference plane, while a second row includes fins forming a negative acute angle A2 with this plane.
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4.发明授权公开(公告)号：US11125091B2
公开(公告)日：2021-09-21
申请号：US16462264
申请日：2017-11-28
发明人： Cédric Zaccardi , Christophe Marcel Lucien Perdrigeon , Mohamed-Lamine Boutaleb , Sébastien Vincent François Dreano
IPC分类号： F01D5/18 , F01D9/06 , F01D25/16 , F02C7/14 , F02C7/18 , F02K3/115 , F28D1/02 , F28F1/40 , F28F7/02 , F28D21/00
摘要： The invention relates to a guide vane for a bypass aircraft turbomachine, its aerodynamic part comprising a first lubricant cooling interior passage in which heat transfer structures are arranged and a second lubricant cooling interior passage in which heat transfer structures are arranged, the aerodynamic part comprising a bent area connecting a lubricant output end of the first interior passage to a lubricant input end of the second passage, the bent area extending along a curved generatrix and being partly delimited by the intrados wall and the extrados wall of the vane. According to the invention, the bent area comprises one or more lubricant guide(s) arranged between the intrados and extrados walls of the vane, and each extending substantially parallel to the curved generatrix of the bent area.
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公开(公告)号：US20180258779A1
公开(公告)日：2018-09-13
申请号：US15914150
申请日：2018-03-07
发明人： Mohamed-Lamine Boutaleb , Fabien Roger Gaston Caty , Sebastien Vincent Francois Dreano , Thierry Georges Paul Papin , Christophe Marcel Lucien Perdrigeon , Cedric Zaccardi
摘要： The invention relates to a guide vane for a twin-spool aircraft turbomachine, of which the aerodynamic part (32) comprises an internal lubricant cooling passage (50a) extending along a principal lubricant flow direction (52a).According to the invention, the aerodynamic part is made in a single piece and also comprises heat transfer fins (80a, 80b) arranged in the passage (50a) connecting the intrados and extrados walls (70, 72) and extending approximately parallel to the direction (52a), these fins being distributed in successive rows along the principal direction (52a) and made such that for two rows (R1, R2) of staggered directly consecutive fins, the row (R1) comprises fins (80a) forming a positive acute angle A1 with a dummy reference plane (Pf), while the row (R2) comprises fins (80b) forming a negative acute angle A2 with this plane (Pf).
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6.发明授权公开(公告)号：US11143045B2
公开(公告)日：2021-10-12
申请号：US15653976
申请日：2017-07-19
发明人： Cédric Zaccardi , Christophe Paul Jacquemard , Thierry Georges Paul Papin , Christophe Marcel Lucien Perdrigeon
IPC分类号： F01D9/06 , F02K3/06 , F02C7/18 , F01D5/18 , F01D9/04 , F01D25/12 , F01D25/18 , F02C7/16 , F28D21/00
摘要： The invention relates to an intermediate case (25) for a twin spool turbomachine for an aircraft, comprising a hub (26), an outer shell (23) and outlet guide vanes (24) installed at their ends on the hub and on the outer shell, and each of at least some of the outlet guide vanes (24) performing a heat exchanger function and comprising a lubricant passage (50a, 50b) designed to be cooled by the fan flow (58) following an outer surface of the outlet guide vane. According to the invention, the case also comprises at least one lubricant duct (55) passing along a circumferential direction of the hub (26) and at least part of which is made from a single casting with the hub, the lubricant duct (55) having at least one lateral opening communicating with the lubricant passage (50a, 50b) of at least one of the vanes (24).
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公开(公告)号：US10830076B2
公开(公告)日：2020-11-10
申请号：US16270771
申请日：2019-02-08
发明人： Cedric Zaccardi , Christophe Marcel Lucien Perdrigeon , Paul Antoine Foresto , Adrien Jacques Philippe Fabre
摘要： A guide vane located in a fan air flow in an aircraft twin-spool turbomachine, the vane being made with an extrados body and an intrados body between which there is a thermal conduction matrix. Furthermore, the attachment devices between the two spools are arranged outside the aerodynamic part of the vane.
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公开(公告)号：US10604235B2
公开(公告)日：2020-03-31
申请号：US15673322
申请日：2017-08-09
发明人： Sébastien Emile Philippe Tajan , Gilles Alain Marie Charier , Clément Cottet , Adrien Jacques Philippe Fabre , Christophe Marcel Lucien Perdrigeon
摘要： The invention relates to turbine engine module (1) including a case (9) rotating around a longitudinal axis (X) and carrying a propeller having a plurality of blades, a stationary case (15) comprising a cylindrical wall (16) extending between an inner wall (17) and an outer wall (18) of the rotating case (9), and a system (26) for changing the pitch of the blades (14) of the propeller. The wall (16) is connected downstream to a first frustoconical wall (42) and upstream to a second frustoconical wall (41), a first rolling bearing (19) being inserted respectively downstream directly between a radially outer face (21) of the inner wall (17) and a radially inner face (23) of the first frustoconical wall, and a second rolling bearing (19′) inserted downstream directly between the radially outer face (21) of the inner wall (17) and an inner face (43) of the second frustoconical wall (41).
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公开(公告)号：US11639665B2
公开(公告)日：2023-05-02
申请号：US17430228
申请日：2020-02-06
摘要： A turbomachine blade including a body that extends mainly in a plane defined by a main axis and a longitudinal direction, which is defined by a lower surface wall, an upper surface wall, a leading edge located at a first longitudinal end of the body and a trailing edge located at a second longitudinal end of the body, wherein the body of the blade includes a plurality of first pipes that extend mainly along the direction of the main axis, for circulation of a gas flow, and a plurality of second pipes that extend mainly along the longitudinal direction, for circulation of a second gas flow.
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公开(公告)号：US20220325753A1
公开(公告)日：2022-10-13
申请号：US17634282
申请日：2020-08-17
发明人： Christophe Marcel Lucien Perdrigeon , Régis Eugène Henri Servant , Guillaume François Jean Bazin
摘要： Devices for distributing oil from a rolling bearing for an aircraft turbine engine include a rolling bearing including two rings, respectively an inner ring and an outer ring, an oil distribution ring configured to be mounted on a turbine engine shaft, said distribution ring including a first outer cylindrical surface for mounting the inner ring of the bearing, an oil recovery scoop supplying a lubricating circuit of the bearing, and an annular track of a dynamic seal. The distribution ring and the track are formed by a single-piece body, and the lubricating circuit is formed in said body and extends into the distribution ring and the track.
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