Patents by Inventor Ba-Phuc TANG
Ba-Phuc TANG has filed for patents to protect the following inventions. This listing includes patent applications that are pending as well as patents that have already been granted by the United States Patent and Trademark Office (USPTO).
-
Patent number: 11867065Abstract: To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.Type: GrantFiled: February 11, 2021Date of Patent: January 9, 2024Assignee: SAFRAN AIRCRAFT ENGINESInventors: Edouard Emmanuel Garreau, Josserand Jacques Andre Bassery, Lucas Geoffrey Desbois, Etienne Leon Francois, Samuel Laurent Noel Mathieu Juge, Elsa Maxime, Ba-Phuc Tang, Denis Gabriel Trahot, Thomas Tsassis
-
Patent number: 11808167Abstract: A turbine engine blade includes an airfoil with a pressure-side wall and a suction-side wall which are connected upstream by a leading edge and downstream by a trailing edge. A cooling circuit has an internal cavity extending inside the airfoil and a plurality of outlet openings, each oriented substantially along a longitudinal axis X. Each outlet opening communicates with the cavity and is arranged in the vicinity of the trailing edge. A calibration device is arranged in the cavity and provided with calibration conduits arranged substantially opposite the outlet openings. The calibration conduits each include an oblong transverse section which is substantially perpendicular to the longitudinal axis.Type: GrantFiled: March 16, 2020Date of Patent: November 7, 2023Assignees: SAFRAN AIRCRAFT ENGINES, SAFRANInventors: Jérémy Jacques Attilio Fanelli, Romain Pierre Cariou, Vianney Simon, Ba-Phuc Tang
-
Publication number: 20230068236Abstract: To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.Type: ApplicationFiled: February 11, 2021Publication date: March 2, 2023Applicant: SAFRAN AIRCRAFT ENGINESInventors: Edouard Emmanuel GARREAU, Josserand Jacques Andre BASSERY, Lucas Geoffrey DESBOIS, Etienne Leon FRANCOIS, Samuel Laurent Noel Mathieu JUGE, Elsa MAXIME, Ba-Phuc TANG, Denis Gabriel TRAHOT, Thomas TSASSIS
-
Patent number: 11499436Abstract: A turbine blade including a root carrying an impeller terminated by a tip in the form of a squealer tip. This impeller also includes a serpentine median circuit, including a first radial pipe collecting air at the root and that is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe, a cavity under the squealer tip running along the pressure side wall, extending from a central region of the tip to the trailing edge, and a radial central pipe collecting air at the root extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip.Type: GrantFiled: November 29, 2019Date of Patent: November 15, 2022Assignees: SAFRAN, SAFRAN AIRCRAFT ENGINESInventors: Léandre Ostino, Pierre Guillaume Auzillon, Romain Pierre Cariou, Thomas Olivier Michel Pierre De Rocquigny, Patrice Eneau, Adrien Bernard Vincent Rollinger, Vianney Simon, Michel Slusarz, Ba-Phuc Tang
-
Patent number: 11421537Abstract: A turbine engine vane including a blade extending following a radial axis and a cooling circuit arranged inside the blade, the cooling circuit including a first cavity and a second cavity arranged downstream from the first cavity following a direction of circulation of a cooling fluid, the first and second cavities extend radially inside the blade and being separated at least partially by a radial partition having a radially external free end which delimits at least partially a passage connecting the first and second cavities, the radial partition connecting a first wall in contact with the outside environment of the blade to a second opposite wall substantially following a transversal axis, perpendicular to the radial axis, respectively in a connection zone. According to the invention, at least one connection zone presents a thickening having a substantially triangular general transversal cross-section.Type: GrantFiled: March 19, 2020Date of Patent: August 23, 2022Assignees: SAFRAN AIRCRAFT ENGINES, SAFRANInventors: Jeremy Jacques Attilio Fanelli, Romain Pierre Cariou, Thomas Olivier Michel Pierre De Rocquigny, Ba-Phuc Tang
-
Patent number: 11396813Abstract: A rough cast blading of this blade includes, a suction sidewall and/or a pressure sidewall of this blading intended to respectively form the suction sidewall and/or the pressure sidewall of the blade, a casting allowance extending over a determined width from a trailing edge of the blading intended to form the trailing edge of the blade in the direction of a leading edge of the blading intended to form the leading edge of the blade, except for a reserved area adjacent to the trailing edge of the blading and whose width is at least one radius of the trailing edge of the blading, over at least part of the height of the blading.Type: GrantFiled: May 23, 2019Date of Patent: July 26, 2022Assignee: Safran Aircraft EnginesInventors: Alexandre Gimel, Josserand Jacques André Bassery, Maxime Paul Numa Givert, Gabriela Mihaila, Marc Soisson, Ba-Phuc Tang
-
Publication number: 20220205365Abstract: A turbine vane includes a root carrying a blade terminated by a squealer tip, the blade having intrados and extrados walls, a leading edge, a trailing edge, and a tip wall delimiting a bottom of the squealer tip, by which the intrados wall is connected to the extrados wall. The blade also includes: a serpentine median circuit, including a first radial pipe that collects air at the root and is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe; a cavity under the squealer tip running along the extrados wall and extending from a central region of the squealer tip to the trailing edge; and a central radial pipe collecting air at the root and extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip.Type: ApplicationFiled: April 24, 2020Publication date: June 30, 2022Inventors: Léandre OSTINO, Pierre Guillaume AUZILLON, Michel SLUSARZ, Patrice ENEAU, Thomas Olivier Michel Pierre DE ROCQUIGNY, Romain Pierre CARIOU, Ba-Phuc TANG, Adrien Bernard Vincent ROLLINGER, Vianney SIMON
-
Publication number: 20220178261Abstract: A turbine engine blade includes an airfoil with a pressure-side wall and a suction-side wall which are connected upstream by a leading edge and downstream by a trailing edge. A cooling circuit has an internal cavity extending inside the airfoil and a plurality of outlet openings, each oriented substantially along a longitudinal axis X. Each outlet opening communicates with the cavity and is arranged in the vicinity of the trailing edge. A calibration device is arranged in the cavity and provided with calibration conduits arranged substantially opposite the outlet openings. The calibration conduits each include an oblong transverse section which is substantially perpendicular to the longitudinal axis.Type: ApplicationFiled: March 16, 2020Publication date: June 9, 2022Applicant: SAFRAN AIRCRAFT ENGINESInventors: Jérémy Jacques Attilio FANELLI, Romain Pierre CARIOU, Vianney SIMON, Ba-Phuc TANG
-
Publication number: 20220025771Abstract: A turbine blade including a root carrying an impeller terminated by a tip in the form of a squealer tip. This impeller also includes a serpentine median circuit, including a first radial pipe collecting air at the root and that is connected by a first bend to a second radial pipe that is connected by a second bend to a third radial pipe, a cavity under the squealer tip running along the pressure side wall, extending from a central region of the tip to the trailing edge, and a radial central pipe collecting air at the root extending between at least two of the three pipes of the median circuit and directly supplying the cavity under the squealer tip.Type: ApplicationFiled: November 29, 2019Publication date: January 27, 2022Applicants: SAFRAN, SAFRAN AIRCRAFT ENGINESInventors: Léandre OSTINO, Pierre Guillaume AUZILLON, Romain Pierre CARIOU, Thomas Olivier Michel Pierre DE ROCQUIGNY, Patrice ENEAU, Adrien Bernard Vincent ROLLINGER, Vianney SIMON, Michel SLUSARZ, Ba-Phuc TANG
-
Publication number: 20210215053Abstract: The invention relates to a movable blade made of aluminum and titanium alloy, for a turbojet engine turbine comprising a vane and at least one root at a distal end of the vane. The root has at least one azimuthal contact surface with another directly adjacent blade. A hard abrasion-resistant material, called wear-resistant material, is deposited onto the at least one azimuthal contact surface. A cavity is produced in said at least one azimuthal contact surface, the wear-resistant material being deposited in the cavity.Type: ApplicationFiled: September 3, 2019Publication date: July 15, 2021Applicant: SAFRAN AIRCRAFT ENGINESInventors: Josserand, Jacques, André BASSERY, Alexandre GIMEL, Ba-Phuc TANG, Vijeay PATEL, Stéphane KNITTEL
-
Publication number: 20210215047Abstract: A rough cast blading of this blade includes, a suction sidewall and/or a pressure sidewall of this blading intended to respectively form the suction sidewall and/or the pressure sidewall of the blade, a casting allowance extending over a determined width from a trailing edge of the blading intended to form the trailing edge of the blade in the direction of a leading edge of the blading intended to form the leading edge of the blade, except for a reserved area adjacent to the trailing edge of the blading and whose width is at least one radius of the trailing edge of the blading, over at least part of the height of the blading.Type: ApplicationFiled: May 23, 2019Publication date: July 15, 2021Applicant: SAFRAN AIRCRAFT ENGINESInventors: Alexandre GIMEL, Josserand Jacques André BASSERY, Maxime Paul Numa GIVERT, Gabriela MIHAILA, Marc SOISSON, Ba-Phuc TANG
-
Patent number: 10787915Abstract: A mobile vane for a turbine engine, including a root designed to be inserted into a receiving element of a rotor disk for a turbine engine, a platform carried by the root, and a blade extending from the platform. The platform includes an upstream edge. The upstream edge includes a lug for engaging in a locking notch of the disk in such a way as to hold the vane axially in relation to the disk, according to the longitudinal direction of the receiving element.Type: GrantFiled: September 28, 2015Date of Patent: September 29, 2020Assignee: SAFRAN AIRCRAFT ENGINESInventors: Jean-Baptise Vincent Desforges, Damien Bernard Quelven, Maurice Guy Judet, Ba-Phuc Tang
-
Publication number: 20200300095Abstract: A turbine engine vane including a blade extending following a radial axis and a cooling circuit arranged inside the blade, the cooling circuit including a first cavity and a second cavity arranged downstream from the first cavity following a direction of circulation of a cooling fluid, the first and second cavities extend radially inside the blade and being separated at least partially by a radial partition having a radially external free end which delimits at least partially a passage connecting the first and second cavities, the radial partition connecting a first wall in contact with the outside environment of the blade to a second opposite wall substantially following a transversal axis, perpendicular to the radial axis, respectively in a connection zone. According to the invention, at least one connection zone presents a thickening having a substantially triangular general transversal cross-section.Type: ApplicationFiled: March 19, 2020Publication date: September 24, 2020Inventors: Jeremy Jacques Attilio FANELLI, Romain Pierre CARIOU, Thomas Olivier Michel Pierre DE ROCQUIGNY, Ba-Phuc TANG
-
Patent number: 10704419Abstract: Turbine distributor sector for an aircraft turbine engine, including an external annular platform sector and an internal annular platform sector, the sectors being coaxial and being connected together by blade assemblies including inner cavities cooled by gas circulation, the external platform sector including through openings of which radially internal ends open into the inner cavities, wherein the external platform sector includes inner ducts for supplying the cavities with gas, the ducts including air outlets opening into the openings and air inlets opening onto a portion of the external annular surface of the external platform sector.Type: GrantFiled: December 5, 2018Date of Patent: July 7, 2020Assignee: SAFRAN AIRCRAFT ENGINESInventors: Josserand Jacques Andre Bassery, Jean-Charles Marcel Bernard Coetard, Raphael Jean Philippe Dupeyre, Etienne Leon Francois, Ba-Phuc Tang
-
Publication number: 20200025037Abstract: Turbine distributor sector for an aircraft turbine engine, including an external annular platform sector and an internal annular platform sector, the sectors being coaxial and being connected together by blade assemblies including inner cavities cooled by gas circulation, the external platform sector including through openings of which radially internal ends open into the inner cavities, wherein the external platform sector includes inner ducts for supplying the cavities with gas, the ducts including air outlets opening into the openings and air inlets opening onto a portion of the external annular surface of the external platform sector.Type: ApplicationFiled: December 5, 2018Publication date: January 23, 2020Inventors: Josserand Jacques Andre BASSERY, Jean-Charles Marcel Bernard COETARD, Raphael Jean Philippe DUPEYRE, Etienne Leon FRANCOIS, Ba-Phuc TANG
-
Patent number: 10280776Abstract: The present invention relates to a turbine assembly (10) of a turbine engine (1), comprising at least: a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) arranged in series, the rotors (12, 14) being mounted on a shaft (2); a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first recess (C1) arranged between the first rotor (12) and the stator (13), from a second recess (C2) arranged between the stator (13) and the second rotor (14); and pressure-reducing means (300, 31) positioned inside the first recess (C1), the assembly being characterized in that said pressure-reducing means (300, 31) comprise a plurality of substantially radial recompression fins (300) extending into the first recess (C1).Type: GrantFiled: December 17, 2015Date of Patent: May 7, 2019Assignee: SAFRAN AIRCRAFT ENGINESInventors: Damien Bernard Quelven, Jean-Baptiste Vincent Desforges, Maurice Guy Judet, Ba-Phuc Tang
-
Patent number: 10196907Abstract: A turbomachine rotor blade includes an outer part at its distal end. The outer part includes a platform defining an outside surface of a passage for gas passing through a turbomachine and presenting first and second opposite side edges; and upstream and downstream sealing wipers extending outwards from the platform, each wiper extending between two lateral faces situated respectively at the first and second side edges. The two lateral faces of the upstream or the downstream wiper are covered at least in part in an anti-wear material.Type: GrantFiled: January 16, 2013Date of Patent: February 5, 2019Assignee: SAFRAN AIRCRAFT ENGINESInventors: Slim Bensalah, Arnaud Negri, Sebastien Digard Brou De Cuissart, Guillaume Klein, Ba-Phuc Tang, David Mathieu, Sibylle Doremus
-
Patent number: 9833834Abstract: A blade preform includes a strut connecting a plat-form to a blade root portion extending longitudinally in an upstream-downstream direction, an upstream web and a downstream web, which each extend in a direction substantially perpendicular to the longitudinal direction of the blade root and are formed at the upstream and downstream ends of the strut. The upstream and downstream webs connect the upstream and downstream ends of the plat-form to the upstream and downstream ends of the blade root. The blade root extends in a direction perpendicular to the longitudinal direction of the blade root over a distance smaller than that of the upstream and downstream webs and the side edges of each web are extended by walls that converge at the flanks of the blade root.Type: GrantFiled: June 4, 2014Date of Patent: December 5, 2017Assignee: SNECMAInventors: Sébastien Serge Francis Congratel, Raphaël Jean Philippe Dupeyre, Guillaume Klein, David Mathieu, Ba-Phuc Tang
-
Patent number: 9827610Abstract: A method for producing a rotor vane (10) for a turbomachine, including producing a rough casting, the heel of which has a downstream lip (121) with a transverse increased thickness (130) such that the lip has an upstream surface (134) substantially parallel to an axis The method further includes machining said increased thickness so that the downstream lip has an upstream surface (138) inclined in relation to the axis.Type: GrantFiled: March 4, 2014Date of Patent: November 28, 2017Assignee: SNECMAInventors: Guillaume Klein, Josserand Bassery, Sebastien Congratel, Raphael Dupeyre, David Mathieu, Ba-Phuc Tang
-
Publication number: 20170328227Abstract: The present invention relates to a turbine assembly (10) of a turbine engine (1), comprising at least: a first bladed rotor (12), a bladed stator (13) and a second bladed rotor (14) arranged in series, the rotors (12, 14) being mounted on a shaft (2); a sealing plate (20) extending between the stator (13) and the shaft (2) and separating a first recess (C1) arranged between the first rotor (12) and the stator (13), from a second recess (C2) arranged between the stator (13) and the second rotor (14); and pressure-reducing means (300, 31) positioned inside the first recess (C1), the assembly being characterised in that said pressure-reducing means (300, 31) comprise a plurality of substantially radial recompression fins (300) extending into the first recess (C1).Type: ApplicationFiled: December 17, 2015Publication date: November 16, 2017Inventors: Damien Bernard QUELVEN, Jean-Baptiste Vincent DESFORGES, Maurice Guy JUDET, Ba-Phuc TANG