determining (1600), from the measured current and voltage, a degree of coking at the dynamic seals.
The invention also relates to a turbine engine comprising a system for monitoring a degree of coking.
The fixed vanes (15) of a diffuser (14) are set to an azimuth angle (a) with respect to the injectors (9) of a combustion chamber so that the paths (35) leading from the trailing edges (34) pass through the gaps (38) between injectors (9), preferably mid-way between same, so that these portions of the flow, which may contain condensed water, do not affect the initiation of combustion.
The invention relates to a turbine engine assembly comprising a housing and a bladed wheel rotatable within the housing. The bladed wheel comprises at least one blade having a head opposite the housing. The assembly is characterized in that the head comprises a magnet and in that the housing comprises a first and second electrical conductor. Each electrical conductor is suitable for generating, across the terminals thereof, an electrical voltage that is induced by the magnet of the head opposite the housing and that represents vibrations to which the head of the blade is subjected when the bladed wheel is rotated. The first electrical conductor includes a first central portion extending around the rotational axis of the bladed wheel and comprising two mutually facing ends, and the second electrical conductor includes a second central portion passing through a space left by the first central portion between the two ends thereof.
G01H 1/00 - Measuring vibrations in solids by using direct conduction to the detector
G01P 3/44 - Devices characterised by the use of electric or magnetic means for measuring angular speed
G01P 3/487 - Devices characterised by the use of electric or magnetic means for measuring angular speed by measuring frequency of generated current or voltage of pulse signals delivered by rotating magnets
4.
ASSEMBLY FOR TURBINE ENGINE FOR MEASURING VIBRATIONS SUSTAINED BY A ROTATING BLADE
The invention relates to an assembly (E) for a turbine engine, the assembly (E) including a casing (1) and an impeller (2) rotatably movable inside the casing (1), the impeller (2) comprising at least one blade (20) having a tip edge (21) opposite the casing (1), the assembly (E) being characterised in that the tip edge (21) comprises a magnet (3) and in that the casing (1) comprises an electrical conductor (4) suitable for generating between the terminals thereof an electric voltage induced by the magnet (3) of the tip edge (21) opposite same and representing vibrations sustained by the tip edge (21) of the blade (20) when the impeller (2) is rotated.
G01H 1/00 - Measuring vibrations in solids by using direct conduction to the detector
G01P 3/44 - Devices characterised by the use of electric or magnetic means for measuring angular speed
G01P 3/487 - Devices characterised by the use of electric or magnetic means for measuring angular speed by measuring frequency of generated current or voltage of pulse signals delivered by rotating magnets
5.
METHOD AND SYSTEM FOR MORE RELIABLE STARTING OF A TURBO MACHINE
The starting system comprises a battery of accumulator cells (110), a DC starter (120), an electronic regulator computer (142), a transmission relay (162), starting accessories (168) and a gas generator (160) itself comprising a compressor (164), a combustion chamber (165) and a high-pressure turbine (166) as well as a free turbine (167). First and second circuits are mounted in parallel and interposed between the battery of accumulator cells (110) and the DC starter (120). The first circuit comprises a DC/DC converter (130) mounted in series with a first circuit breaker (132) and the second circuit comprises a second circuit breaker (133). Furthermore, the system comprises at least one sensor (163) sensing the rotational speed of the compressor (164), a sensor (151) sensing the temperature at the inlet of the free turbine (167), and a control circuit (141) controlling the first and second circuit breakers (132, 133) on the basis of the information supplied by the sensor (163) sensing the rotational speed of the compressor (164) and by the sensor (151) sensing the temperature at the inlet to the free turbine (167).
The invention relates to an assembly (E) for a turbine engine, the assembly (E) including a casing (1) and an impeller (2) rotatably movable inside the casing (1), the impeller (2) comprising at least one blade (20) having a tip edge (21) opposite the casing (1), the assembly (E) being characterised in that the tip edge (21) comprises a magnet (3) and in that the casing (1) comprises an electrical conductor (4) suitable for generating between the terminals thereof an electric voltage induced by the magnet (3) of the tip edge (21) opposite same and representing vibrations sustained by the tip edge (21) of the blade (20) when the impeller (2) is rotated.
G01H 1/00 - Measuring vibrations in solids by using direct conduction to the detector
G01P 3/44 - Devices characterised by the use of electric or magnetic means for measuring angular speed
G01P 3/487 - Devices characterised by the use of electric or magnetic means for measuring angular speed by measuring frequency of generated current or voltage of pulse signals delivered by rotating magnets
7.
METHOD AND SYSTEM FOR MORE RELIABLE STARTING OF A TURBO MACHINE
The starting system comprises a battery of accumulator cells (110), a DC starter (120), an electronic regulator computer (142), a transmission relay (162), starting accessories (168) and a gas generator (160) itself comprising a compressor (164), a combustion chamber (165) and a high-pressure turbine (166) as well as a free turbine (167). First and second circuits are mounted in parallel and interposed between the battery of accumulator cells (110) and the DC starter (120). The first circuit comprises a DC/DC converter (130) mounted in series with a first circuit breaker (132) and the second circuit comprises a second circuit breaker (133). Furthermore, the system comprises at least one sensor (163) sensing the rotational speed of the compressor (164), a sensor (151) sensing the temperature at the inlet of the free turbine (167), and a control circuit (141) controlling the first and second circuit breakers (132, 133) on the basis of the information supplied by the sensor (163) sensing the rotational speed of the compressor (164) and by the sensor (151) sensing the temperature at the inlet to the free turbine (167).
The invention relates to a turbine engine assembly comprising a housing and a bladed wheel rotatable within the housing. The bladed wheel comprises at least one blade having a head opposite the housing. The assembly is characterized in that the head comprises a magnet and in that the housing comprises a first and second electrical conductor. Each electrical conductor is suitable for generating, across the terminals thereof, an electrical voltage that is induced by the magnet of the head opposite the housing and that represents vibrations to which the head of the blade is subjected when the bladed wheel is rotated. The first electrical conductor includes a first central portion extending around the rotational axis of the bladed wheel and comprising two mutually facing ends, and the second electrical conductor includes a second central portion passing through a space left by the first central portion between the two ends thereof.
G01H 1/00 - Measuring vibrations in solids by using direct conduction to the detector
G01P 3/44 - Devices characterised by the use of electric or magnetic means for measuring angular speed
G01P 3/487 - Devices characterised by the use of electric or magnetic means for measuring angular speed by measuring frequency of generated current or voltage of pulse signals delivered by rotating magnets
The invention relates to a fluid-draining device (10) for an aircraft engine, including a main drain (11) configured such as to collect fluids drained from the engine, characterised in that said device comprises means (13, 19) for pumping the fluids contained in the main drain and discharging said fluids, and supervision means (14) configured to signal an abnormal collection of fluids by the main drain, said surveillance means being configured such as to activate when the flow of collected fluids is greater than the flow from the pumping means.
The invention relates to a method for anticorrosion and antiwear treatment of an aluminium or aluminium alloy substrate, especially in the field of aeronautics, for protecting mechanical parts of aeroplanes or helicopters subjected to both corrosion and wear. According to the invention, the method comprises applying to the substrate (10): a step of sol-gel treatment forming a sol-gel layer (200); and, following the sol-gel treatment step, a step of hard oxidation forming a hard oxide layer (30).
The invention relates to the field of testing mechanical parts, and in particular to an endoscope (1) capable of being used for frequency testing of a part which is difficult to access, as well as to a method for using said endoscope (i), which includes an endoscopic head (2), a device (3) for displaying images picked up via said endoscopic head (2), and an elongate member (4) connecting the endoscopic head (2) to the display device (3), and in which the endoscopic head (2) also includes a frequency testing device (7) which includes at least one vibration sensor (10) intended for picking up a vibration response of a subject of frequency testing.
G02B 23/24 - Instruments for viewing the inside of hollow bodies, e.g. fibrescopes
G01N 29/14 - Investigating or analysing materials by the use of ultrasonic, sonic or infrasonic wavesVisualisation of the interior of objects by transmitting ultrasonic or sonic waves through the object using acoustic emission techniques
12.
METHOD FOR AUTOMATICALLY CONTROLLING THE OPERATING SPEED OF A HELICOPTER TURBOSHAFT ENGINE, CORRESPONDING CONTROL DEVICE AND HELICOPTER PROVIDED WITH SUCH A DEVICE
The invention relates to a method for automatically controlling an operating speed of a helicopter turboshaft engine, including a step (10) of receiving data (27, 28, 29) that represent the flight of the helicopter; a step (11) of selecting the turboshaft engine for which a change in speed would be more relevant; a step (12) of determining an operating speed of said turboshaft engine, referred to as chosen speed, selected among a plurality of predetermined operating speeds; and a step (14) of controlling the operating speed of said turboshaft engine to said chosen speed. The invention also relates to a corresponding control device.
The invention relates to the field of testing mechanical parts, and in particular to an endoscope (1) capable of being used for frequency testing of a part which is difficult to access, as well as to a method for using said endoscope (i), which includes an endoscopic head (2), a device (3) for displaying images picked up via said endoscopic head (2), and an elongate member (4) connecting the endoscopic head (2) to the display device (3), and in which the endoscopic head (2) also includes a frequency testing device (7) which includes at least one vibration sensor (10) intended for picking up a vibration response of a subject of frequency testing.
G02B 23/24 - Instruments for viewing the inside of hollow bodies, e.g. fibrescopes
G01N 29/14 - Investigating or analysing materials by the use of ultrasonic, sonic or infrasonic wavesVisualisation of the interior of objects by transmitting ultrasonic or sonic waves through the object using acoustic emission techniques
The invention relates to a fluid-draining device (10) for an aircraft engine, including a main drain (11) configured such as to collect fluids drained from the engine, characterised in that said device comprises means (13, 19) for pumping the fluids contained in the main drain and discharging said fluids, and supervision means (14) configured to signal an abnormal collection of fluids by the main drain, said surveillance means being configured such as to activate when the flow of collected fluids is greater than the flow from the pumping means.
F02C 7/232 - Fuel valvesDraining valves or systems
15.
METHOD FOR AUTOMATICALLY CONTROLLING THE OPERATING SPEED OF A HELICOPTER TURBOSHAFT ENGINE, CORRESPONDING CONTROL DEVICE AND HELICOPTER PROVIDED WITH SUCH A DEVICE
The invention relates to a method for automatically controlling an operating speed of a helicopter turboshaft engine, including a step (10) of receiving data (27, 28, 29) that represent the flight of the helicopter; a step (11) of selecting the turboshaft engine for which a change in speed would be more relevant; a step (12) of determining an operating speed of said turboshaft engine, referred to as chosen speed, selected among a plurality of predetermined operating speeds; and a step (14) of controlling the operating speed of said turboshaft engine to said chosen speed. The invention also relates to a corresponding control device.
The invention relates to a method for anticorrosion and antiwear treatment of an aluminium or aluminium alloy substrate, especially in the field of aeronautics, for protecting mechanical parts of aeroplanes or helicopters subjected to both corrosion and wear. According to the invention, the method comprises applying to the substrate (10): a step of sol-gel treatment forming a sol-gel layer (200); and, following the sol-gel treatment step, a step of hard oxidation forming a hard oxide layer (30).
The invention relates to a bladed rotor including a rotor disc (12) having two front surfaces (14, 15) and an outer peripheral surface (16), cavities (18) being arranged in the outer peripheral surface (16) and in communication with at least one of the front surfaces (14, 15). The rotor (10) includes blades (30) each having a root (32) by which the blade is secured in a cavity (18), an end surface (31) of the root being substantially level with the front surface (14) of the disc when the blade is secured in the cavity. A coating layer (40) is deposited on the disk (12) so as to cover both at least one portion of the front surface (14) of the disc and at least one portion of the end surface (31) of the root (32).
The invention relates to a bladed rotor including a rotor disc (12) having two front surfaces (14, 15) and an outer peripheral surface (16), cavities (18) being arranged in the outer peripheral surface (16) and in communication with at least one of the front surfaces (14, 15). The rotor (10) includes blades (30) each having a root (32) by which the blade is secured in a cavity (18), an end surface (31) of the root being substantially level with the front surface (14) of the disc when the blade is secured in the cavity. A coating layer (40) is deposited on the disk (12) so as to cover both at least one portion of the front surface (14) of the disc and at least one portion of the end surface (31) of the root (32).
The invention relates to a turbine engine combustion assembly, including: a casing; a fuel combustion chamber arranged in the casing, the combustion chamber including two rotationally symmetrical walls, internal and external, respectively, one extending inside the other and being connected by a chamber bottom wall; and at least one prevaporization tube extending in the combustion chamber from the external wall thereof, the prevaporization tube being mounted on the casing and penetrating the combustion chamber through an opening provided in the external wall thereof, characterized in that the casing includes a ferrule provided with an opening facing the opening of the external wall of the combustion chamber, and a tubular embossing added to said ferrule around the opening, the prevaporization tube being inserted in the embossing and extending through the opening of the casing ferrule, and in that the assembly further includes a fuel injector leading into the prevaporization tube. The prevaporization tube includes a base mounted on the embossing, the fuel injector includes a plate mounted on the base of the prevaporization tube, and an injection pipe extends into the tube. The invention further relates to a turbine engine including such an assembly.
F23R 3/32 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices being tubular
The invention relates to a method for controlling a turbine engine (5) which includes a compressor (8), a combustion chamber (9), first and second turbines (10, 12), a first rotary shaft (11) rotatably secured to said compressor and to said first turbine, a second rotary shaft (13) to which the second turbine is rotatably secured, said second rotary shaft being freely rotatable relative to the first rotary shaft, and a controller (15) for controlling the fuel supply to the combustion chamber. Said controller interrupts the fuel supply to the combustion chamber if a speed of rotation (N2) of said second rotary shaft exceeds a maximum threshold (N2imax) which varies according to at least one indicative physical parameter associated with a mechanical power extracted from combustion gases in the second turbine.
F01D 21/02 - Shutting-down responsive to overspeed
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
The invention relates to a method for controlling a turbine engine (5) which includes a compressor (8), a combustion chamber (9), first and second turbines (10, 12), a first rotary shaft (11) rotatably secured to said compressor and to said first turbine, a second rotary shaft (13) to which the second turbine is rotatably secured, said second rotary shaft being freely rotatable relative to the first rotary shaft, and a controller (15) for controlling the fuel supply to the combustion chamber. Said controller interrupts the fuel supply to the combustion chamber if a speed of rotation (N2) of said second rotary shaft exceeds a maximum threshold (N2imax) which varies according to at least one indicative physical parameter associated with a mechanical power extracted from combustion gases in the second turbine.
F01D 21/02 - Shutting-down responsive to overspeed
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
The invention relates to a method for optimising the specific consumption of a helicopter provided with two turboshafts (1, 2), each one comprising a gas generator (11, 21) provided with a combustion chamber (CC), each of said turboshafts (1, 2) being able to operate alone in the continuous flight mode, the other turboshaft (2, 1) being then in a so-called super-slow mode with zero power and with the combustion chamber lit, said super-slow mode being assisted by a mechanical rotation of the shaft (AE) of the gas generator in this mode, in such a way as to reduce the operating temperature and the fuel consumption of said gas generator.
The invention concerns the field of turbomachines, and more particularly a turbomachine (5a) comprising at least a compressor (8), a combustion chamber (9), a first turbine (10) linked to the compressor (8) by a first rotating shaft (11), a device (13) for actuating said first rotating shaft (11) in order to keep the first turbine (10) and the compressor (8) rotating when the combustion chamber (9) is off, and a lubrication circuit (14) for lubricating the turbomachine (5a). This circuit (14) passes through at least one heat source (15, 15a-15c) capable of heating a lubricant in said lubrication circuit (14) when the first turbine (10) and the compressor (8) are rotating while the combustion chamber (9) is off.
The invention concerns the field of turbomachines, and more particularly a turbomachine (5a) comprising at least a compressor (8), a combustion chamber (9), a first turbine (10) linked to the compressor (8) by a first rotating shaft (11), a device (13) for actuating said first rotating shaft (11) in order to maintain the rotation of the first turbine (10) and the compressor (8) when the combustion chamber (9) is off, and a lubrication circuit (14) for lubricating the turbomachine (5a). This circuit (14) passes through at least one heat source (15, 15a-15c) capable of heating a lubricant in said lubrication circuit (14) when the first turbine (10) and the compressor (8) are rotating while the combustion chamber (9) is off.
The invention concerns the field of turbomachines, and more particularly a turbomachine (5a) comprising at least a compressor (8), a combustion chamber (9), a first turbine (10) linked to the compressor (8) by a first rotating shaft (11), a device (13) for actuating said first rotating shaft (11) in order to keep the first turbine (10) and the compressor (8) rotating when the combustion chamber (9) is off, and a lubrication circuit (14) for lubricating the turbomachine (5a). This circuit (14) passes through at least one heat source (15, 15a-15c) capable of heating a lubricant in said lubrication circuit (14) when the first turbine (10) and the compressor (8) are rotating while the combustion chamber (9) is off.
Two channel high energy generator. The invention relates to an ignition device for an aircraft engine comprising at least two spark plugs, the device comprising a power supply, a first channel (3) for supplying power to a first spark plug and a second channel (4) for supplying power to a second spark plug, said channels (3, 4) being linked with the power supply by supply distribution means (5) controlled by an FADEC type control system, wherein said distribution means (5) comprises a first circuit (6) configured to alternately supply said first channel (3) or said second channel (4), and a second circuit (7) for simultaneously supplying said first (3) and second channels (4), the device being configured to use either the first circuit (6) or the second circuit (7) during a start-up.
Two channel high energy generator. The invention relates to an ignition device for an aircraft engine comprising at least two spark plugs, the device comprising a power supply, a first channel (3) for supplying power to a first spark plug and a second channel (4) for supplying power to a second spark plug, said channels (3, 4) being linked with the power supply by supply distribution means (5) controlled by an FADEC type control system, wherein said distribution means (5) comprises a first circuit (6) configured to alternately supply said first channel (3) or said second channel (4), and a second circuit (7) for simultaneously supplying said first (3) and second channels (4), the device being configured to use either the first circuit (6) or the second circuit (7) during a start-up.
The invention concerns the field of turbomachines, and more particularly a turbomachine (5a) comprising at least a compressor (8), a combustion chamber (9), a first turbine (10) linked to the compressor (8) by a first rotating shaft (11), a device (13) for actuating said first rotating shaft (11) in order to maintain the rotation of the first turbine (10) and the compressor (8) when the combustion chamber (9) is off, and a lubrication circuit (14) for lubricating the turbomachine (5a). This circuit (14) passes through at least one heat source (15, 15a-15c) capable of heating a lubricant in said lubrication circuit (14) when the first turbine (10) and the compressor (8) are rotating while the combustion chamber (9) is off.
System (20) for the emergency starting of aircraft turbomachines, comprising at least one solid-fuel gas generator (22), an electrically operated ignition device (24), a computer (28) connected to the ignition device, and at least two independent starters (18) each one intended to start a turbomachine, each starter comprising a turbine (38) for driving a shaft (34) intended to be coupled to a shaft (54) of the corresponding turbomachine, the gas outlet of the generator being connected to the inlet (44) of the turbine of each starter by one and the same distribution valve (26) connected to the computer (28).
The invention relates to a method (1000) for monitoring a degree of coking at dynamic seals of a turbine engine comprising: a gas generator including a rotary shaft and an injection wheel mounted on the shaft; a fuel injection manifold designed to convey fuel to the injection wheel; and dynamic seals designed to provide a seal between the injection wheel and the fuel injection manifold. The method is characterised in that it comprises steps consisting in: measuring (1100) the rotation speed of the gas generator shaft during a turbine engine autorotation phase, and determining (1200) a degree of coking at the dynamic seals, based on changes in the measured rotation speed over time. The invention also relates to a turbine engine equipped with a monitoring system.
F02C 3/16 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant the combustion chambers being formed at least partly in the turbine rotor
F23R 3/38 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising rotary fuel injection means
31.
MONITORING OF A DEGREE OF COKING AT DYNAMIC SEALS BY A STARTER MOTOR
The invention relates to a method (1000) for monitoring a degree of coking at dynamic seals of a turbine engine comprising: a gas generator including a rotary shaft and an injection wheel mounted on the shaft; a fuel injection manifold; dynamic seals for providing a seal between the injection wheel and the fuel injection manifold; and a starter motor. The method comprises steps consisting in: during a phase in which the rotation of the gas generator shaft is initiated by the starter motor, measuring (1500) a current flowing through the starter motor and a voltage at the terminals of the starter motor, and determining (1600) a degree of coking at the dynamic seals, based on the current and voltage measured. The invention also relates to a turbine engine equipped with a system for monitoring the degree of coking.
The invention relates to a method (1000) for monitoring a degree of coking at dynamic seals of a turbine engine comprising: a gas generator including a rotary shaft and an injection wheel mounted on the shaft; a fuel injection manifold designed to convey fuel to the injection wheel; and dynamic seals designed to provide a seal between the injection wheel and the fuel injection manifold. The method is characterised in that it comprises steps consisting in: measuring (1100) the rotation speed of the gas generator shaft during a turbine engine autorotation phase, and determining (1200) a degree of coking at the dynamic seals, based on changes in the measured rotation speed over time. The invention also relates to a turbine engine equipped with a monitoring system.
F02C 3/16 - Gas-turbine plants characterised by the use of combustion products as the working fluid characterised by the arrangement of the combustion chamber in the plant the combustion chambers being formed at least partly in the turbine rotor
F23R 3/38 - Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising rotary fuel injection means
33.
MONITORING OF A DEGREE OF COKING AT DYNAMIC SEALS BY A STARTER MOTOR
The invention relates to a method (1000) for monitoring a degree of coking at dynamic seals of a turbine engine comprising: a gas generator including a rotary shaft and an injection wheel mounted on the shaft; a fuel injection manifold; dynamic seals for providing a seal between the injection wheel and the fuel injection manifold; and a starter motor. The method comprises steps consisting in: during a phase in which the rotation of the gas generator shaft is initiated by the starter motor, measuring (1500) a current flowing through the starter motor and a voltage at the terminals of the starter motor, and determining (1600) a degree of coking at the dynamic seals, based on the current and voltage measured. The invention also relates to a turbine engine equipped with a system for monitoring the degree of coking.
The invention consists of a hybrid cutoff member (100; 500) for an electric circuit comprising a static cutoff component (101; 501) and an electromechanical cutoff component, characterised in that the static component (101; 501) is fixed on a support (110; 510) bearing electrical contacts (111, 112; 511, 512) for the static component, said support (110; 510) being configured to move, on receiving a cutoff command, so as to withdraw at least one of said electrical contacts (111, 112; 511, 512) from the respective pin of same, thus forming said electromechanical cutoff component.
The invention consists of a hybrid cutoff member (100; 500) for an electric circuit comprising a static cutoff component (101; 501) and an electromechanical cutoff component, characterised in that the static component (101; 501) is fixed on a support (110; 510) bearing electrical contacts (111, 112; 511, 512) for the static component, said support (110; 510) being configured to move, on receiving a cutoff command, so as to withdraw at least one of said electrical contacts (111, 112; 511, 512) from the respective pin of same, thus forming said electromechanical cutoff component.
The present invention relates to a turbomachine comprising a casing (7) with an internal wall (3i) that forms a wall of an air duct (3) and at least one opening (7r) passing through the casing, opening into said duct (3) and forming a passage for an endoscope, the opening (7r), while the turbomachine is running, being plugged by a plug (8) having an end surface portion (8s) in the continuation of the internal wall (3i), characterized in that a wear indicator indicative of wear of the internal wall of the casing is associated with the plug (8) or with the internal wall (3i) of the casing in the vicinity of the plug (8). The means of the invention allows inspection that is easy and does not require the use of a measurement instrument.
The present invention relates to a turbomachine comprising a casing (7) with an internal wall (3i) that forms a wall of an air duct (3) and at least one opening (7r) passing through the casing, opening into said duct (3) and forming a passage for an endoscope, the opening (7r), while the turbomachine is running, being plugged by a plug (8) having an end surface portion (8s) in the continuation of the internal wall (3i), characterized in that a wear indicator indicative of wear of the internal wall of the casing is associated with the plug (8) or with the internal wall (3i) of the casing in the vicinity of the plug (8). The means of the invention allows inspection that is easy and does not require the use of a measurement instrument.
Test rig (100) for oligocyclic fatigue and possibly combined oligocyclic and polycylic fatigue testing, to reproduce a bearer for turbomachine components, comprising a support member (126) fixed to a structure (108) and defining at least one bearing surface (148), and a test specimen (110) which is connected to traction means (104) in order to force the test specimen to bear against the or each bearing surface of the member, the or each bearing surface (148) being borne by an element (136) which is mounted such that it can rotate about a first axis (B) on the support member, and the test specimen being connected to the traction means by means (114) of articulation about a second axis (A) substantially perpendicular to the first axis, the rig further comprising means for adjusting and immobilising the element and the test specimen in positions around the aforementioned axes.
Test rig (100) for oligocyclic fatigue and possibly combined oligocyclic and polycylic fatigue testing, to reproduce a bearer for turbomachine components, comprising a support member (126) fixed to a structure (108) and defining at least one bearing surface (148), and a test specimen (110) which is connected to traction means (104) in order to force the test specimen to bear against the or each bearing surface of the member, the or each bearing surface (148) being borne by an element (136) which is mounted such that it can rotate about a first axis (B) on the support member, and the test specimen being connected to the traction means by means (114) of articulation about a second axis (A) substantially perpendicular to the first axis, the rig further comprising means for adjusting and immobilising the element and the test specimen in positions around the aforementioned axes.
A test bench (100) combining high-frequency tribological stress and oligocyclic fatigue, comprising a first test piece (114) fixed to a frame (108) and defining at least one bearing surface, a second test piece (110) that is linked to traction means for exerting bearing stress on the second test piece on the or each bearing surface of the first test piece, means (126) for heating the test pieces and means (122) for exerting vibrational stress on the test pieces to carry out a fretting fatigue and oligocyclic and polycyclic fatigue test, one of the test pieces comprising a portion (112) having the shape of a blade root of a turbomachine rotor and being engaged in a groove (116) having a shape that is substantially complementary to the other of the test pieces in order to reproduce a turbomachine blade/disc fastener. A test bench (100) combining high-frequency tribological stress and oligocyclic fatigue, comprising a first test piece (114) fixed to a frame (108) and defining at least one bearing surface, a second test piece (110) that is linked to traction means for exerting bearing stress on the second test piece on the or each bearing surface of the first test piece, means (126) for heating the test pieces and means (122) for exerting vibrational stress on the test pieces to carry out a fretting fatigue and oligocyclic and polycyclic fatigue test, one of the test pieces comprising a portion (112) having the shape of a root of a turbomachine rotor blade and being engaged in a groove (116) having a shape that is substantially complementary to the other of the test pieces in order to reproduce a turbomachine blade/disc fastener.
The invention relates to a turbomachine air intake casing (10) comprising an inner annular wall (12) and an outer annular wall (14) delimiting an air passage (16), and at least two nozzles for injecting cleaning agent, in which a first nozzle (18) is oriented towards the outer wall (14) while a second nozzle (20) is oriented towards the inner wall (12).
The invention relates to a vane (10) of a diffuser (5) for a radial or mixed-flow compressor (2) of an engine (1), including a leading edge (11) arranged facing a flow of gas, a trailing edge (12) being opposite the leading edge (11), a side upper surface wall (13) and a side lower surface wall (14) which connect the leading edge (11) to the trailing edge (12), and a profile including a curved line (15) having at least two points of inflection (I1, I2) between the leading edge (11) and the trailing edge (12). The invention also relates to a corresponding radial diffuser (2).
The invention relates to a vane (10) of a diffuser (5) for a radial or mixed-flow compressor (2) of an engine (1), including a leading edge (11) arranged facing a flow of gas, a trailing edge (12) being opposite the leading edge (11), a side upper surface wall (13) and a side lower surface wall (14) which connect the leading edge (11) to the trailing edge (12), and a profile including a curved line (15) having at least two points of inflection (I1, I2) between the leading edge (11) and the trailing edge (12). The invention also relates to a corresponding radial diffuser (2).
A turbine ring for a turbomachine, in particular for a helicopter, of which the vibratory behaviour is reduced. According to the invention, this turbine ring comprises an essentially cylindrical support (31), and one or a plurality of sectors (32) forming a crown configured to provide an air flow section, each sector (32) being fixed to the support (31) by an anchoring device (33a, 33b), in which the anchoring device (33a) comprises an anchoring part (35) belonging to the support (31) and protruding towards the sector (32), and an anchoring part (34) belonging to the sector (32) and protruding towards the support (31), the anchoring parts of the support (34) and of the sector (35) being configured to engage in order to fasten the sector (32) to the support (31), the ring further comprising a damping device (50) provided within the anchoring device (33a) and radially constrained between a portion of the sector (34e) and a portion of the support (31i) so as to dampen the relative movements of the sector (32) in relation to the support (31).
A dosing device for an engine fuel supply circuit comprising a dosing valve (21), and a pressure control device keeping a constant pressure differential between the downstream and upstream of the dosing valve (21), in which the dosing valve (21) comprises a seat (31), provided with an inlet opening (31e) and an outlet opening (31s), a shutter (32), disposed within the seat (31), and an actuator, controlling the position of the shutter (32), and in which the shutter (32) defines a passage between the inlet opening (31e) and the outlet opening (31s) of which the minimum cross-section is variable on the basis of the position of the shutter (32) along a path extending between a lower stop and an upper stop and passing through a threshold position. According to the invention, the shutter (32) is configured in such a way that the minimum cross-section of said passage, and therefore the flow of fuel passing through the valve (21), increases linearly on the basis of the coordinate of the position of the shutter between the lower stop and the threshold position, and that the minimum cross-section of said passage, and therefore the flow of fuel, increases quadratically, or more quickly, on the basis of the coordinate of the position of the shutter between the threshold position and the upper stop.
A turbine ring for a turbomachine, in particular for a helicopter, of which the vibratory behaviour is reduced. According to the invention, this turbine ring comprises an essentially cylindrical support (31), and one or a plurality of sectors (32) forming a crown configured to provide an air flow section, each sector (32) being fixed to the support (31) by an anchoring device (33a, 33b), in which the anchoring device (33a) comprises an anchoring part (35) belonging to the support (31) and protruding towards the sector (32), and an anchoring part (34) belonging to the sector (32) and protruding towards the support (31), the anchoring parts of the support (34) and of the sector (35) being configured to engage in order to fasten the sector (32) to the support (31), the ring further comprising a damping device (50) provided within the anchoring device (33a) and radially constrained between a portion of the sector (34e) and a portion of the support (31i) so as to dampen the relative movements of the sector (32) in relation to the support (31).
A method for monitoring the health of an aircraft (10) turbomachine (12), comprising a step (E3) of detecting a period from measurements of at least one first parameter relative to the turbomachine (12) and of at least one second parameter relative to the flight conditions of said aircraft (10), to the aircraft equipment (10) and/or to the turbomachine (12), recorded over time during a normal flight of said aircraft (10), referred to as a stable period, during which at least the second parameter is substantially stable for a predefined length of time, at least the measurements of the first parameter during said stable period being intended to be analysed in order to determine the operating state of the turbomachine.
G01M 15/02 - Details or accessories of testing apparatus
G01M 15/14 - Testing gas-turbine engines or jet-propulsion engines
B64F 5/00 - Designing, manufacturing, assembling, cleaning, maintaining or repairing aircraft, not otherwise provided forHandling, transporting, testing or inspecting aircraft components, not otherwise provided for
48.
METHOD FOR MONITORING A DEGREE OF CLOGGING OF THE STARTING INJECTORS OF A TURBINE ENGINE
The invention relates to a method for monitoring a degree of clogging of the starting injectors of a turbine engine, which includes: a combustion chamber into which at least one starting injector supplied with fuel leads, said starting injectors being suitable for initiating the combustion in said chamber by igniting the fuel; and a turbine rotated by the gases resulting from the combustion of the fuel in the chamber, the method being characterised in that it includes the steps involving: the measurement (1100), during a phase of starting the turbine engine, of the temperature of the exhaust gases at the outlet of the turbine; and the determination (1200), from the changes over time in the temperature thus measured, of a degree of clogging of the starting injectors. The invention also provides a system for monitoring a degree of clogging capable of implementing the method, and a turbine engine including such a system.
The invention relates to a method for monitoring a degree of clogging of the starting injectors of a turbine engine, which includes: a combustion chamber into which at least one starting injector supplied with fuel leads, said starting injectors being suitable for initiating the combustion in said chamber by igniting the fuel; and a turbine rotated by the gases resulting from the combustion of the fuel in the chamber, the method being characterised in that it includes the steps involving: the measurement (1100), during a phase of starting the turbine engine, of the temperature of the exhaust gases at the outlet of the turbine; and the determination (1200), from the changes over time in the temperature thus measured, of a degree of clogging of the starting injectors. The invention also provides a system for monitoring a degree of clogging capable of implementing the method, and a turbine engine including such a system.
The invention relates to the broaching of at least one slot (3) in a part such as a turbine rotor disk (4) or a turbomachine compressor disk, said slot (3) being machined by means of a broach (1) inclined at a broaching angle (a). Said broach (1) has an inter-tooth pitch (P) that is a sub-multiple of the length to be broached (L).
The invention relates to the broaching of at least one slot (3) in a part such as a turbine rotor disk (4) or a turbomachine compressor disk, said slot (3) being machined by means of a broach (1) inclined at a broaching angle (a). Said broach (1) has an inter-tooth pitch (P) that is a sub-multiple of the length to be broached (L).
Turbo machine combustion assembly (1) comprising a combustion chamber (10), at least one starting injector (17), a plurality of main injectors (18) distributed at constant angular intervals around the circumference of the combustion chamber, each starting injector being positioned between two consecutive main injectors, equal distances therefrom, and a fuel supply circuit (40) supplying fuel to the injectors, in which assembly the combustion chamber is delimited by two axisymmetric walls - an external wall (14) and an internal wall (12) - which are connected by an annular chamber end wall (16), the fuel supply circuit being designed to supply at least one starting injector continuously, each continuously-supplied starting injector being oriented toward the chamber end wall and dimensioned to spread a spray (F) of fuel between 120° and 180° wide, and the flow rate of fuel injected by the main injectors (18') between which the starting injectors are positioned being reduced by comparison with the flow rate injected by the other main injectors (18).
Turbo machine combustion assembly (1) comprising a combustion chamber (10), at least one starting injector (17), a plurality of main injectors (18) distributed at constant angular intervals around the circumference of the combustion chamber, each starting injector being positioned between two consecutive main injectors, equal distances therefrom, and a fuel supply circuit (40) supplying fuel to the injectors, in which assembly the combustion chamber is delimited by two axisymmetric walls - an external wall (14) and an internal wall (12) - which are connected by an annular chamber end wall (16), the fuel supply circuit being designed to supply at least one starting injector continuously, each continuously-supplied starting injector being oriented toward the chamber end wall and dimensioned to spread a spray (F) of fuel between 120° and 180° wide, and the flow rate of fuel injected by the main injectors (18') between which the starting injectors are positioned being reduced by comparison with the flow rate injected by the other main injectors (18).
The invention relates to a system for recommending maintenance of helicopter engines depending on the technical condition of the engine, the standard replacement of parts between engines, and the replacement of parts with different parts. The system comprises: a centralized database storing data relating to (i) working condition and working condition indicators, (ii) modifications of the engines, (iii) maintenance plans for the engines, (iv) causes of unscheduled events, (v) maintenance applied to the engines, and (vi) instantiated configurations; means for acquiring the working condition indicators and for updating the working condition data; means for identifying maintenance to be applied to the engines depending on the data; means for generating an alarm for identified maintenance to be performed; means for the digitally-signed updating of the applied maintenance and configuration data according to maintenance operations; and means for deactivating an alarm once the maintenance associated with the alarm is completed.
The invention relates to a turbomachine comprising an air compression stage (2) comprising at least one moving compressor wheel (20), an air intake pipe (4) coupled to said air compression stage (2), a first sealing device (54) positioned between a front portion (56) of the moving compressor wheel (20) and the air intake pipe (4) comprising at least one pressure seal (60), a channel (45) for conveying air compressed by the moving wheel (20) and a second sealing device (64) positioned between a rear portion (66) of the moving compressor wheel (20) and the conveyance channel (45) and configured to receive an air flow (F2) coming from the conveyance channel (45), the turbomachine being remarkable in that the second sealing device (64) is configured to allow a withdrawal (F3) of part of the air passing through it and in that the withdrawn air (F3) is conveyed to the pressure seal (60) of the first sealing device (54) so as to keep it pressurized.
The invention relates to the field of turbomachines and more specifically to a device (13) and to a method for temporarily increasing the power of at least one first turbomachine (5A). This device (13) comprises a tank (14) of liquid coolant and a first injection circuit (16A) connected to said tank (14) and opening onto at least one injection nozzle (22) that can be installed upstream of at least one compressor stage (8) of the first turbomachine (5A). This first injection circuit (16A) comprises at least a first switchover valve (23) configured to open when an overpressure crosses a predetermined threshold with respect to a pressure downstream of at least one compressor stage (8) of a second turbomachine (5B) so as to allow the liquid coolant to flow to the said injection nozzle (22) of the first injection circuit (16A).
The invention relates to the field of turbomachines and more specifically to a device (13) and to a method for temporarily increasing the power of at least one first turbomachine (5A). This device (13) comprises a tank (14) of liquid coolant and a first injection circuit (16A) connected to said tank (14) and opening onto at least one injection nozzle (22) that can be installed upstream of at least one compressor stage (8) of the first turbomachine (5A). This first injection circuit (16A) comprises at least a first switchover valve (23) configured to open when an overpressure crosses a predetermined threshold with respect to a pressure downstream of at least one compressor stage (8) of a second turbomachine (5B) so as to allow the liquid coolant to flow to the said injection nozzle (22) of the first injection circuit (16A).
The invention relates to a turbomachine comprising an air compression stage (2) comprising at least one moving compressor wheel (20), an air intake pipe (4) coupled to said air compression stage (2), a first sealing device (54) positioned between a front portion (56) of the moving compressor wheel (20) and the air intake pipe (4) comprising at least one pressure seal (60), a channel (45) for conveying air compressed by the moving wheel (20) and a second sealing device (64) positioned between a rear portion (66) of the moving compressor wheel (20) and the conveyance channel (45) and configured to receive an air flow (F2) coming from the conveyance channel (45), the turbomachine being remarkable in that the second sealing device (64) is configured to allow a withdrawal (F3) of part of the air passing through it and in that the withdrawn air (F3) is conveyed to the pressure seal (60) of the first sealing device (54) so as to keep it pressurized.
The invention relates to the field of laser welding, and in particular to a laser welding head (1) intended to be fixed under a lens for focusing the laser, the laser welding head (1) comprising at least one annular nozzle (5) for the injection of a shielding gas, and a chamber (3) for protecting the focusing lens with a transverse current of air. The annular nozzle (5) is arranged around a clear optical axis (0) passing through the laser welding head (1). The chamber (3) for protecting the focusing lens with a transverse current of air comprises an air inlet (9) and an air exhaust (10) square with the air inlet (9) in a plane substantially perpendicular to the said optical axis (0). This laser welding head (1) is configured to be fixed against said focusing lens with no lateral aperture between the focusing lens and said protecting chamber (3). It ensures a distance (d) of at least 100 mm between an outlet (18) of the annular nozzle (5) and said protecting chamber (3).
B23K 26/14 - Working by laser beam, e.g. welding, cutting or boring using a fluid stream, e.g. a jet of gas, in conjunction with the laser beamNozzles therefor
The invention relates to the field of laser welding, and in particular to a laser welding head (1) intended to be fixed under a lens for focusing the laser, the laser welding head (1) comprising at least one annular nozzle (5) for the injection of a shielding gas, and a chamber (3) for protecting the focusing lens with a transverse current of air. The annular nozzle (5) is arranged around a clear optical axis (0) passing through the laser welding head (1). The chamber (3) for protecting the focusing lens with a transverse current of air comprises an air inlet (9) and an air exhaust (10) square with the air inlet (9) in a plane substantially perpendicular to the said optical axis (0). This laser welding head (1) is configured to be fixed against said focusing lens with no lateral aperture between the focusing lens and said protecting chamber (3). It ensures a distance (d) of at least 100 mm between an outlet (18) of the annular nozzle (5) and said protecting chamber (3).
B23K 26/14 - Working by laser beam, e.g. welding, cutting or boring using a fluid stream, e.g. a jet of gas, in conjunction with the laser beamNozzles therefor
The invention relates to an air conditioning system for the pressurized cabin of an aircraft. The system (1) comprises an air intake module (3) that is designed to draw in ambient air external to the aircraft, an air compression module (5) that is designed to compress the drawn-in air flow (F1), and an air cooling module (10) that is designed to cool the compressed air flow (F2, F3) from a cryogenic fluid, said cooling module (10) comprising a condenser (12) for condensing the water in the air flow, a water extractor (13) for extracting said water, a cooler (14) for cooling the dry air flow originating from the water extractor (13) and a tank (15) of a cryogenic fluid by means of which the water from the air flow is condensed in the condenser (12) and the dry air resulting from the extractor is cooled in the cooler (14).
B64D 13/08 - Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space the air being conditioned the air being heated or cooled
B64D 13/06 - Arrangements or adaptations of air-treatment apparatus for aircraft crew or passengers, or freight space the air being conditioned
The invention relates to a compression assembly for a turbine engine, in particular a turboshaft engine, said assembly including an air-intake duct capable of receiving an air flow, at least one air-compression stage including at least one mobile compressor impeller (115) onto which the duct leads, and a pre-rotary vane (105) positioned in the air-intake duct upstream from the mobile compressor impeller (115) in order to control the speed of the air of said flow at the intake of the mobile impeller and including a plurality of variable-pitch blades (110), the assembly being characterised in that the pitch (S2) between two consecutive blades (110) of the vane (105) is greater than the chord (C2) of one of the two blades (110) at a given height of the air duct, preferably in the upper portion thereof.
The invention concerns a rotor-stator assembly for a gas-turbine engine, comprising a rotor (2) at the apex of which a layer (8) made from a ceramic material forming an abrasive coating is deposited, said layer consisting mainly of zirconia and having a level of porosity equal to or lower than 15%, and a stator (4) disposed around the rotor and provided, opposite the rotor apex, with a layer (6) made from a ceramic material forming an abradable coating, said layer consisting mainly of zirconia and having a level of porosity of between 20% and 50%, with pores smaller than or equal to 50 µm.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
C04B 35/48 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on oxides based on zirconium or hafnium oxides or zirconates or hafnates
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
The invention concerns a rotor-stator assembly for a gas-turbine engine, comprising a rotor (2) at the apex of which a layer (8) made from a ceramic material forming an abrasive coating is deposited, said layer consisting mainly of zirconia and having a level of porosity equal to or lower than 15%, and a stator (4) disposed around the rotor and provided, opposite the rotor apex, with a layer (6) made from a ceramic material forming an abradable coating, said layer consisting mainly of zirconia and having a level of porosity of between 20% and 50%, with pores smaller than or equal to 50 μm.
F01D 11/12 - Preventing or minimising internal leakage of working fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible, deformable or resiliently biased part
F01D 5/28 - Selecting particular materialsMeasures against erosion or corrosion
C04B 35/48 - Shaped ceramic products characterised by their compositionCeramic compositionsProcessing powders of inorganic compounds preparatory to the manufacturing of ceramic products based on oxides based on zirconium or hafnium oxides or zirconates or hafnates
The invention concerns a device (1) for a turbomachine chamber injector, characterised in that it comprises: a conduit (2) having at least: a first opening (3), and a second opening (4), a blocking system (5) for blocking the conduit, making it possible to regulate the passage of fluids into the conduit (2), the blocking system (5) being configured to: allow fuel to flow from the first opening (3) to the second opening (4) only from a first pressure (P1) of said fuel, for the passage of a supply of fuel, and allow air to flow from the second opening (4) to the first opening (3) only from a second pressure (P2) of said air, for the passage of purge air, the second pressure being greater than the first pressure.
The invention relates to a turbine engine combustion assembly comprising a casing, a combustion chamber, and at least one fuel injector for starting a turbine engine, said combustion chamber being defined by two walls of revolution, namely an internal wall and an external wall extending one inside the other and being connected by an annular chamber base wall. The external wall of the chamber is secured to an annular external wall of the casing. The injector is attached to the annular external wall of the casing and comprises a fuel ignition enclosure extending inside the casing successively through an opening in the casing wall and an opening in the external wall of the combustion chamber before opening into said chamber. At least one wall of the ignition enclosure that extends between the casing wall and the combustion chamber wall is provided with at least one air intake port. The combustion assembly is characterised in that the external wall of the combustion chamber is solidly connected to a device for plugging the air intake port(s) according to the thermal expansion state of the combustion chamber.
The invention relates to a start-up injector for a turbine engine combustion chamber, said injector comprising: a fuel injection circuit; and a fuel ignition circuit including a fuel injector supplied by the fuel injection circuit and a spark plug (101) for igniting the injected fuel. The start-up injector is characterised in that it also comprises: a partitioned enclosure including a first compartment (106) in which the fuel is ignited by the spark plug (101) and a second compartment (107) separated from the first compartment by a thermally conductive partition (105); and a main combustion start-up circuit which includes at least one fuel injector supplied by the fuel injection circuit and opens into the second compartment (107) of the enclosure such as to inject the fuel against the wall (105). The invention also relates to a combustion assembly and a turbine engine comprising at least one such start injector.
The invention concerns a combustion chamber (1), comprising -an annular outer housing (10), -the flame tube (20), arranged inside the outer housing (10) of the combustion chamber (1), and -fixing means (30) for fixing the wall (22) of the flame tube (20) to the outer housing (10) of the combustion chamber (1), comprising exactly three centring members (30a) distributed in a fixing plane (P) normal to the longitudinal axis (X) of the flame tube (20), at least one of the fixing means (30) being an injector or a spark plug.
The invention concerns a test bench for a sealing system, which uses a machining apparatus, particularly a high speed machining apparatus. At least one test piece forming a housing is mounted on the workpiece support, while a labyrinth or other seal test piece is mounted on the tool-holder part, which makes it possible to control the rotation and the movement of the labyrinth test piece in the housing test piece.
G01M 3/00 - Investigating fluid tightness of structures
G01M 3/28 - Investigating fluid tightness of structures by using fluid or vacuum by measuring rate of loss or gain of fluid, e.g. by pressure-responsive devices, by flow detectors for pipes, cables, or tubesInvestigating fluid tightness of structures by using fluid or vacuum by measuring rate of loss or gain of fluid, e.g. by pressure-responsive devices, by flow detectors for pipe joints or sealsInvestigating fluid tightness of structures by using fluid or vacuum by measuring rate of loss or gain of fluid, e.g. by pressure-responsive devices, by flow detectors for valves
G01M 3/38 - Investigating fluid tightness of structures by using light
F16J 15/34 - Sealings between relatively-moving surfaces with slip-ring pressed against a more or less radial face on one member
The invention concerns a device (1) for a turbomachine chamber injector, characterised in that it comprises: a conduit (2) having at least: a first opening (3), and a second opening (4), a blocking system (5) for blocking the conduit, making it possible to regulate the passage of fluids into the conduit (2), the blocking system (5) being configured to: allow fuel to flow from the first opening (3) to the second opening (4) only from a first pressure (P1) of said fuel, for the passage of a supply of fuel, and allow air to flow from the second opening (4) to the first opening (3) only from a second pressure (P2) of said air, for the passage of purge air, the second pressure being greater than the first pressure.
The aim of the present invention is to allow, in a reduced axial space, a free expansion of the envelope of a centrifugal compressor cover in the radial direction during increases in temperature, in order to manage the control of the radial clearances of the leading edge and the axial clearances of the trailing edge. For this purpose, the invention provides a cover (30) in which the junction (41) between the fastener (4) and the envelope (3) is formed between substantially the middle (I) of the envelope (3) and the trailing edge (BF) of same and in which, when cold, the fastener (4) has a shape (43) that is essentially concave and elastically deformable by bending. This concave shape (43) is connected to the envelope (3) by a wall (5) that extends at the junction (41) by means of a convex shape (44), such that the fastener (4) has, overall, a cross-sectional configuration in the shape of an "S" with a junction (41) positioned closer to the trailing edge (BF) of the envelope (3) than to the leading edge (BA) of same.
F01D 11/18 - Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
F04D 29/42 - CasingsConnections for working fluid for radial or helico-centrifugal pumps
The invention relates to a turbine engine combustion assembly comprising a casing, a combustion chamber, and at least one fuel injector for starting a turbine engine, said combustion chamber being defined by two walls of revolution, namely an internal wall and an external wall extending one inside the other and being connected by an annular chamber base wall. The external wall of the chamber is secured to an annular external wall of the casing. The injector is attached to the annular external wall of the casing and comprises a fuel ignition enclosure extending inside the casing successively through an opening in the casing wall and an opening in the external wall of the combustion chamber before opening into said chamber. At least one wall of the ignition enclosure that extends between the casing wall and the combustion chamber wall is provided with at least one air intake port. The combustion assembly is characterised in that the external wall of the combustion chamber is solidly connected to a device for plugging the air intake port(s) according to the thermal expansion state of the combustion chamber.
The invention relates to a start-up injector for a turbine engine combustion chamber, said injector comprising: a fuel injection circuit; and a fuel ignition circuit including a fuel injector supplied by the fuel injection circuit and a spark plug (101) for igniting the injected fuel. The start-up injector is characterised in that it also comprises: a partitioned enclosure including a first compartment (106) in which the fuel is ignited by the spark plug (101) and a second compartment (107) separated from the first compartment by a thermally conductive partition (105); and a main combustion start-up circuit which includes at least one fuel injector supplied by the fuel injection circuit and opens into the second compartment (107) of the enclosure such as to inject the fuel against the wall (105). The invention also relates to a combustion assembly and a turbine engine comprising at least one such start injector.
The invention relates to an accessory drive gearbox for an aircraft turbine engine, said gearbox (140) comprising a box (42), a linkage (115) for controlling flaps of the aircraft, which linkage is designed to slide axially inside said gearbox (140), and an actuator (120) for driving said linkage (115) and mounted on said box (42), said actuator (120) comprising a hollow body (121), a piston (123) designed for translational movement inside said body (121) and a piston rod (112) connected to said piston (123) and extending at least partially outside the body (121) of the actuator (120), said rod (122) being connected to the linkage (115), the gearbox (140) being characterized in that the body (121) of the actuator (120) is positioned between the connection between the rod (112) and the linkage (115), and the box (42) of the gearbox (140).
The invention relates to a method of starting a turboengine for an aircraft, said turboengine comprising a combustion chamber, a compressor shaft on which is mounted a compressor wheel for feeding compressed air to said combustion chamber, at least one starter linked to said shaft in such a way as to provide it with the starting torque of determined value sufficient to drive it in rotation. The method comprises a step (E1) of acceleration of the compressor shaft during a first starting phase, then a step (E2) of stabilizing the speed of rotation of the compressor shaft during a second starting phase. During the acceleration step (E1), the speed of rotation of the shaft is regulated in such a way that the acceleration of the shaft remains substantially constant.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE NORMALE SUPERIEURE DE CACHAN (France)
Inventor
Mary, Caroline
Cluzel, Christophe
De Moura Pinho, Raul, Fernando
Longuet, Arnaud
Pommier, Sylvie
Vogel, François
Abstract
A high frequency method for determining the non-propagation threshold of fatigue cracks in which a cyclic load (32, 32A) is applied to at least one test piece comprising, in a test area (10A), an elliptical hole (12) having a notch (14) at one apex, and held between two rigid masses (24, 26), two rigid prestressing plates (20, 22) being disposed on either side of said test piece and being secured at each of the two ends (20A, 22A; 20B, 22B) of same to the two rigid masses, the frequency of the cyclic load (32, 32A) being chosen as equal to the natural frequency of the test piece/masses/prestressing plates assembly so as to generate a fatigue crack from the notch; then, when it is observed that the crack has stopped propagating, the final length of the crack is recorded, and said non-propagation threshold ΔKth of the fatigue crack is determined using a table, the cyclic load being obtained by a vibrating electrodynamic pot integrally attached by means of rigid posts to a frame supporting the two rigid masses and comprising a push rod to transmit said cyclic load to the test piece/masses/prestressing plates assembly.
G01N 3/38 - Investigating strength properties of solid materials by application of mechanical stress by applying repeated or pulsating forces generated by electromagnetic means
G01N 3/32 - Investigating strength properties of solid materials by application of mechanical stress by applying repeated or pulsating forces
The invention relates to a method of starting a turboengine for an aircraft, said turboengine comprising a combustion chamber, a compressor shaft on which is mounted a compressor wheel for feeding compressed air to said combustion chamber, at least one starter linked to said shaft in such a way as to provide it with the starting torque of determined value sufficient to drive it in rotation. The method comprises a step (E1) of acceleration of the compressor shaft during a first starting phase, then a step (E2) of stabilizing the speed of rotation of the compressor shaft during a second starting phase. During the acceleration step (E1), the speed of rotation of the shaft is regulated in such a way that the acceleration of the shaft remains substantially constant.
CENTRE NATIONAL DE LA RECHERCHE SCIENTIFIQUE (France)
ECOLE NORMALE SUPERIEURE DE CACHAN (France)
Inventor
Mary, Caroline
Cluzel, Christophe
De Moura Pinho, Raul Fernando
Longuet, Arnaud
Pommier, Sylvie
Vogel, Francois
Abstract
A high frequency method for determining the non-propagation threshold of fatigue cracks in which a cyclic load (32, 32A) is applied to at least one test piece comprising, in a test area (10A), an elliptical hole (12) having a notch (14) at one apex, and held between two rigid masses (24, 26), two rigid prestressing plates (20, 22) being disposed on either side of said test piece and being secured at each of the two ends (20A, 22A; 20B, 22B) of same to the two rigid masses, the frequency of the cyclic load (32, 32A) being chosen as equal to the natural frequency of the test piece/masses/prestressing plates assembly so as to generate a fatigue crack from the notch; then, when it is observed that the crack has stopped propagating, the final length of the crack is recorded, and said non-propagation threshold ?Kth of the fatigue crack is determined using a table, the cyclic load being obtained by a vibrating electrodynamic pot integrally attached by means of rigid posts to a frame supporting the two rigid masses and comprising a push rod to transmit said cyclic load to the test piece/masses/prestressing plates assembly.
G01N 3/38 - Investigating strength properties of solid materials by application of mechanical stress by applying repeated or pulsating forces generated by electromagnetic means
The invention relates to an accessory drive gearbox for an aircraft turbine engine, said gearbox (140) comprising a box (42), a linkage (115) for controlling flaps of the aircraft, which linkage is designed to slide axially inside said gearbox (140), and an actuator (120) for driving said linkage (115) and mounted on said box (42), said actuator (120) comprising a hollow body (121), a piston (123) designed for translational movement inside said body (121) and a piston rod (112) connected to said piston (123) and extending at least partially outside the body (121) of the actuator (120), said rod (122) being connected to the linkage (115), the gearbox (140) being characterized in that the body (121) of the actuator (120) is positioned between the connection between the rod (112) and the linkage (115), and the box (42) of the gearbox (140).
The aim of the invention is to prevent the formation of smoke at least during the restarting of the engines. To this end, the fuel purged during the engine shutdown is trapped and drained during the beginning of the following flight, making use of the incline of the helicopter. A collector (4) of a helicopter engine comprises an outer longitudinal wall (41) and two closed end walls (42, 43), a longitudinal axis of symmetry (X'X) inclined in an ascending manner, couplings (51 to 53) to be connected to the purge drains, and a connection (54) coupled to a gas jet nozzle (5) and connected to the bottom end wall (43). The collector (4) also comprises, in the inner space (V) thereof, an enclosure (6) having an axis of symmetry (E'E) substantially parallel to the axis of the collector (X'X). The enclosure (6) has a longitudinal wall (61) and two transverse end walls (62, 63). The enclosure (6) is connected to the purge coupling of the injection wheel (53) via a radial connection (64) joining the longitudinal wall (61) thereof, the axis of symmetry of the enclosure (E'E) being inclined in relation to the horizontal ground of reference (S0) when the helicopter is in the ground position (H0), by an angle of reference (A0) such that said axis (E'E) is parallel to the ground of reference (S0) when the helicopter is in the acceleration phase.
The aim of the invention is to prevent the formation of smoke at least during the restarting of the engines. To this end, the fuel purged during the engine shutdown is trapped and drained during the beginning of the following flight, making use of the incline of the helicopter. A collector (4) of a helicopter engine comprises an outer longitudinal wall (41) and two closed end walls (42, 43), a longitudinal axis of symmetry (X'X) inclined in an ascending manner, couplings (51 to 53) to be connected to the purge drains, and a connection (54) coupled to a gas jet nozzle (5) and connected to the bottom end wall (43). The collector (4) also comprises, in the inner space (V) thereof, an enclosure (6) having an axis of symmetry (E'E) substantially parallel to the axis of the collector (X'X). The enclosure (6) has a longitudinal wall (61) and two transverse end walls (62, 63). The enclosure (6) is connected to the purge coupling of the injection wheel (53) via a radial connection (64) joining the longitudinal wall (61) thereof, the axis of symmetry of the enclosure (E'E) being inclined in relation to the horizontal ground of reference (S0) when the helicopter is in the ground position (H0), by an angle of reference (A0) such that said axis (E'E) is parallel to the ground of reference (S0) when the helicopter is in the acceleration phase.
The invention relates to the field of turbine engine injection tubes and more specifically to a turbine engine injection tube (101) including: a first set of transfer tubes (102) connected so as to form a main circuit for supplying fuel to at least a first and second injector set (100, 1000); and a second set of transfer tubes (102) connected parallel to the first set so as to form an auxiliary circuit for supplying fuel to said first injector set (100). The tube (101) also includes in particular at least one dual connection (104) having at least one first tip (501), in which one end of a transfer tube (102) from the main circuit is received; one second tip (502), in which one end of a transfer tube (502) from the auxiliary circuit is received; and a mounting surface having a first opening, fluidly connected to the first tip (501), and a second opening (512), fluidly connected to the second tip (502). Said surface for mounting the dual connection (104) is capable of connecting the dual connection (104) to an injector of said first injector set (100).
The invention relates to the field of turbine engine injection tubes and more specifically to a turbine engine injection tube (101) including: a first set of transfer tubes (102) connected so as to form a main circuit for supplying fuel to at least a first and second injector set (100, 1000); and a second set of transfer tubes (102) connected parallel to the first set so as to form an auxiliary circuit for supplying fuel to said first injector set (100). The tube (101) also includes in particular at least one dual connection (104) having at least one first tip (501), in which one end of a transfer tube (102) from the main circuit is received; one second tip (502), in which one end of a transfer tube (502) from the auxiliary circuit is received; and a mounting surface having a first opening, fluidly connected to the first tip (501), and a second opening (512), fluidly connected to the second tip (502). Said surface for mounting the dual connection (104) is capable of connecting the dual connection (104) to an injector of said first injector set (100).
Helicopter engine air intake provided with an anti-icing grating that offers significant bypass flow in the event of icing. According to the invention, this air intake comprises air intake lips (30, 32), and an anti-icing grating (36) mounted on the external ends (30a, 32a) of the air intake lips (30, 32) and which is interposed in the flow of air entering the air intake (34), at least one air intake lip (30, 32) being formed of thin sheet metal.
B64D 33/02 - Arrangement in aircraft of power plant parts or auxiliaries not otherwise provided for of combustion air intakes
F02C 7/055 - Air intakes for gas-turbine plants or jet-propulsion plants having provisions for obviating the penetration of damaging objects or particles with intake grids, screens or guards
Filling device (100, 200, 300, 400, 500) for fluid tank (10) comprising a filling pipe (102, 402, 502), a first stopper (110, 210, 420, 510) to prevent the overfilling of the tank, and a second stopper (120, 220, 320, 420, 520) for preventing unwanted discharge of fluid from the tank; a first float (110, 210, 410, 510) mechanically connected to the first stopper so that the first float, by positioning itself in a predetermined position, closes the first stopper; and a system (122, 222, 422) for retaining the second stopper which, when fluid leaves the tank, closes the second stopper and, when the fluid enters the tank, opens the second stopper. The retaining system constantly urges the second stopper because when the device is in a position under the effect of the weight of a heavy element (430, 520), the retaining system tends constantly to keep the second stopper (420, 520) in its closed position.
The aim of the invention is to optimize the entirety of the drive power available in a helicopter by using an auxiliary motor to supply power to the equipment and accessories of the helicopter that are connected to the engines. In an example of an optimized power transfer architecture for implementing the invention, the main engines (1, 1') and the APU group (8), as an auxiliary motor, comprise a gas generator (2; 81) connected, for the main engines (1, 1'), to the gearboxes (6) and accessory boxes (7) of mechanical, electric, and/or hydraulic power sockets, and connected, for the APU group (8), to at least one power conversion member (83, 84, 11). The power conversion member (83, 84, 11) of the APU group (8) is connected to the equipment and accessories via the gearbox (6) and/or via the accessory box (7) of the main engines (1, 1').
B64D 35/08 - Transmitting power from power plants to propellers or rotorsArrangements of transmissions characterised by the transmission being driven by a plurality of power plants
87.
METHOD AND CONFIGURATION FOR AN AUXILIARY POWER ENGINE TO DELIVER PROPULSIVE AND/OR NON-PROPULSIVE ENERGY IN A HELICOPTER ARCHITECTURE
The invention aims to optimise all of the engine power available on a helicopter provided with an auxiliary engine by allowing that auxiliary engine to provide non-propulsive and/or propulsive energy during flight. For this purpose, this auxiliary engine is coupled in such a way as to be able to directly contribute to the supply of propulsive mechanical or electrical energy and non-propulsive electrical energy of the aircraft. An example configuration architecture comprises an onboard electrical network (2), two main engines (5a, 5b) and an energy conversion system for converting mechanical/electrical energy (6, 6a, 6b, 7) between a main gearbox BTP (40) transferring power to the propulsion members (4, 41) and means for receiving electrical energy comprising the onboard network (2) and a power electronics system (9) linked with starters (8) of the main engines (5a, 5b). The configuration also comprises an auxiliary power engine (3) supplying electrical energy to the means for receiving electrical energy (2, 9) via the energy conversion system (6, 6a, 6b, 7) and mechanical coupling means (8a, 11a to 11d) between the auxiliary engine (3, 0, 30) and at least one propulsion means (4, 41).
The invention seeks to provide emergency starters that allow responsiveness of this order of magnitude, namely within a few seconds, without having the disadvantages associated with the mass and size of the backup hydraulic or pneumatic starters mentioned hereinabove. To achieve this, the present invention proposes coupling an instantaneous gas thrust of pyrotechnic type with a positive displacement transmission generator in conjunction with automatic coupling to/uncoupling from the set that is to be started. An emergency start-up system (10) according to the invention comprises at least one pyrotechnic gas generator (5) connected to an electric initiator (3) itself connected to a computer, a positive displacement motor (100) housing straight-cut spur gears, the pyrotechnic gas generator (5) being coupled to the motor (100) by an inlet (121) in the casing (120). The motor (100) comprises a means of connection capable of moving at one end of the transmission shaft (40b) so as to be able to couple this transmission shaft to a receiving shaft of the set that is to be started via a centrifugal clutch (170).
The aim of the invention is to prevent the backflow of hot primary air into the peripheral opening formed between the nozzle and the ejector of the exhaust stream of a gas turbine. For this purpose, according to the invention, said peripheral opening is partially closed to prevent the backflow of the primary flow into the engine bay. More specifically, the invention comprises a method for discharging exhaust gas from a gas turbine wherein the number of sectors, the position and the angle at the centre (C) of at least one sector (21) of the peripheral opening (1) capable of forming an area for the re-intake of the primary flow (Fp) into the engine bay (Mb) are determined by correlation of the interactions between the secondary cooling flows (Fs) and the primary flow (Fp), from the following behaviour parameters: air gyration and speed at the inlet of the nozzle (2), geometry of the exhaust stream (2, 3), routing of the secondary cooling flow (Fs) for cooling the engine bay (Mb), and the geometry and position of the inlets (E1) of the secondary flows (Fs). Said peripheral opening is then closed on the angular sector(s) (21) identified in this way.
The aim of the invention is to prevent the backflow of hot primary air into the peripheral opening formed between the nozzle and the ejector of the exhaust stream of a gas turbine. For this purpose, according to the invention, said peripheral opening is partially closed to prevent the backflow of the primary flow into the engine bay. More specifically, the invention comprises a method for discharging exhaust gas from a gas turbine wherein the number of sectors, the position and the angle at the centre (C) of at least one sector (21) of the peripheral opening (1) capable of forming an area for the re-intake of the primary flow (Fp) into the engine bay (Mb) are determined by correlation of the interactions between the secondary cooling flows (Fs) and the primary flow (Fp), from the following behaviour parameters: air gyration and speed at the inlet of the nozzle (2), geometry of the exhaust stream (2, 3), routing of the secondary cooling flow (Fs) for cooling the engine bay (Mb), and the geometry and position of the inlets (E1) of the secondary flows (Fs). Said peripheral opening is then closed on the angular sector(s) (21) identified in this way.
The invention relates to the field of turbine engines, and more particularly, to a bracket (14) for at least one bearing of the hot section of a turbine engine, comprising at least one central hub (15) including an outer bearing seat for directly receiving the bearing (13), an annular casing segment (16) around the central hub (15), and a plurality of radial arms (17) connecting said central hub (15) to said annular segment (16). The radial arms (17) are inclined in an axial direction and in a tangential direction, and built into the central hub (15) and annular casing segment (16) so as to form an integral part.
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
The invention relates to the field of turbine engines, and more particularly, to a bracket (14) for at least one bearing of the hot section of a turbine engine, comprising at least one central hub (15) including an outer bearing seat for directly receiving the bearing (13), an annular casing segment (16) around the central hub (15), and a plurality of radial arms (17) connecting said central hub (15) to said annular segment (16). The radial arms (17) are inclined in an axial direction and in a tangential direction, and built into the central hub (15) and annular casing segment (16) so as to form an integral part.
F01D 25/16 - Arrangement of bearingsSupporting or mounting bearings in casings
F02C 3/10 - Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with another turbine driving an output shaft but not driving the compressor
93.
TURBINE ENGINE COMPRISING AN ELECTRICALLY ACTIVATED FUEL SUPPLY PUMP, AND TURBINE ENGINE FUEL SUPPLY METHOD
The invention relates to an aircraft turbine engine comprising a turbine engine shaft and a pumping module (100) which includes a structural casing (9), a pumping shaft (11) connected to the turbine engine shaft, a turbine engine fuel supply pump (3) that is mounted on said pumping shaft (11) and is mounted inside the structural casing, and an electrical device (5) that is mounted on the pumping shaft (11) and is suitable for rotatably driving the pumping shaft (11) to actuate the supply pump (3) or for being rotatably driven by the pumping shaft (11) to supply electrical power to a device (8) of the turbine engine. The electrical device comprises rotor elements (51) mounted at the outer periphery of a movable part (32) of the supply pump as well as stator elements (52) mounted at the inner periphery of the structural casing.
The invention relates to an aircraft turbine engine comprising a turbine engine shaft and a pumping module (100) which includes a structural casing (9), a pumping shaft (11) connected to the turbine engine shaft, a turbine engine fuel supply pump (3) that is mounted on said pumping shaft (11) and is mounted inside the structural casing, and an electrical device (5) that is mounted on the pumping shaft (11) and is suitable for rotatably driving the pumping shaft (11) to actuate the supply pump (3) or for being rotatably driven by the pumping shaft (11) to supply electrical power to a device (8) of the turbine engine. The electrical device comprises rotor elements (51) mounted at the outer periphery of a movable part (32) of the supply pump as well as stator elements (52) mounted at the inner periphery of the structural casing.
With a view to optimizing the resources needed to continue the flight upon the loss of at least one main engine, the invention seeks to unburden the propulsion system that has remained operative of all or some of the non-propulsive energy tappings through a continuously-operating additional generation of non-propulsive power. More specifically the method involves, using the power production unit GPP, constantly in operation during flight and taking over some (PGA) of the nominal total non-propulsive power (PTA) of the aircraft, to supply, almost instantaneously, increased non-propulsive powers (PSU, PMU, PIU) according to at least three respective emergency settings (RS, RM, RI) upon said engine failure. The GPP control and monitoring function assesses the time elapsed for each emergency setting (RS, RM, RI), and informs the data processing unit of this time with an alarm being emitted if the operating times (Xmax, Ymax, Zmax) allocated to each emergency setting (RS, RM, RI) are exceeded, and the emergency function adjusts the aircraft's non-propulsive power tappings between the main engines and the unit GPP either automatically or on the orders of the pilot.
With a view to optimizing the resources needed to continue the flight upon the loss of at least one main engine, the invention seeks to unburden the propulsion system that has remained operative of all or some of the non-propulsive energy tappings through a continuously-operating additional generation of non-propulsive power. More specifically the method involves, using the power production unit GPP, constantly in operation during flight and taking over some (PGA) of the nominal total non-propulsive power (PTA) of the aircraft, to supply, almost instantaneously, increased non-propulsive powers (PSU, PMU, PIU) according to at least three respective emergency settings (RS, RM, RI) upon said engine failure. The GPP control and monitoring function assesses the time elapsed for each emergency setting (RS, RM, RI), and informs the data processing unit of this time with an alarm being emitted if the operating times (Xmax, Ymax, Zmax) allocated to each emergency setting (RS, RM, RI) are exceeded, and the emergency function adjusts the aircraft's non-propulsive power tappings between the main engines and the unit GPP either automatically or on the orders of the pilot.
A control device for controlling pivotable vanes of a turbo-machine, including: a plurality of pivotable vanes distributed in azimuth over at least 90° around the axis of the turbo-machine, the pivotable vanes being oriented substantially radially relative to the axis of the turbo-machine; and a control ring portion for controlling pivoting of the vanes, each vane being connected to the control ring portion by a link, the control ring portion being held around the axis of the turbo-machine by the links; wherein at least two of the links are connected to the ring portion by respective ball-joint connections, with other links being connected to the ring portion via respective sliding pivot connections.
The invention relates to the optimised recovery of energy in an aircraft, at altitude and on the ground, using a single architecture. To this end, the invention aims to recover thermal energy from the exhaust. An architecture for the recovery of energy comprises an auxiliary power unit APU (20) fitted with an exhaust nozzle (14) and a gas generator (2a) fitted with a shaft (21) for transmitting power to a load compressor (22). Said compressor supplies compressed air via a supply duct (C1) to the ECS air conditioning system (30) of the passenger cabin (40). Additionally, a recovery turbocharger (10) is connected, directly or via a transmission case, to the shaft (21) of the APU unit (20). Said turbocharger (10) comprises a recovery turbine (11) powered by a downstream branch (C3b) of a conduit (C3) fitted to a heat exchanger (1) mounted on the nozzle (14). Said conduit (C3) has an upstream branch (C3a) connected to channels (41, 42) connecting the air outlets of the cabin (40) and the compressor (12). A second exchanger (2) can be fitted between the supply duct (C1) and the cabin outlet channel (41).
The invention relates to the optimised recovery of energy in an aircraft, at altitude and on the ground, using a single architecture. To this end, the invention aims to recover thermal energy from the exhaust. An architecture for the recovery of energy comprises an auxiliary power unit APU (20) fitted with an exhaust nozzle (14) and a gas generator (2a) fitted with a shaft (21) for transmitting power to a load compressor (22). Said compressor supplies compressed air via a supply duct (C1) to the ECS air conditioning system (30) of the passenger cabin (40). Additionally, a recovery turbocharger (10) is connected, directly or via a transmission case, to the shaft (21) of the APU unit (20). Said turbocharger (10) comprises a recovery turbine (11) powered by a downstream branch (C3b) of a conduit (C3) fitted to a heat exchanger (1) mounted on the nozzle (14). Said conduit (C3) has an upstream branch (C3a) connected to channels (41, 42) connecting the air outlets of the cabin (40) and the compressor (12). A second exchanger (2) can be fitted between the supply duct (C1) and the cabin outlet channel (41).
An annular wall of a combustion chamber (10) of a turbo engine, comprising a cold side (16a, 18a) and a hot side (16b, 18b), a plurality of primary and dilution holes (30) distributed in a circumferential row to allow air circulating on the cold side (16a, 18a) of the annular wall to penetrate into the hot side (16b, 18b) in order provide the dilution of an air/fuel mixture; and a plurality of cooling holes (32) to allow air circulating on the cold side (16a, 18a) of the annular wall to penetrate into the hot side (16b, 18b) in order to form a film of cooling air along the annular wall, the cooling holes being distributed in a plurality of circumferential rows spaced axially apart from one another, and the geometrical axes of each of the cooling holes being inclined, in an axial direction of flow D of the combustion gases, by an angle of inclination A1 relative to a normal N of the annular wall. The wall further comprises a plurality of additional cooling holes (34) arranged directly downstream from the dilution holes and distributed in a plurality of circumferential rows spaced axially apart from one another, the geometrical axes of each of the additional cooling holes being arranged in a plane perpendicular to said axial direction D and inclined by an angle of inclination 82 relative to a normal N of said annular wall.
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