Abstract: The invention relates to a pre-vaporizing pipe (5) for a combustion chamber of a turbomachine, comprising a generally elongate main body (50), comprising a first end portion (51) delimiting an inlet duct (52) in which a fuel injector is to be mounted, the body (50) having an external surface (500) having a first length (L1) that extends from the portion (51) to at least one first fuel outlet end piece (53), and a second length (L2) that extends the first length (L1) from the end piece (53) and that extends from the end piece (53) to a second fuel outlet end piece (54). According to the invention, inside the body (50) there is a first channel (55) connecting the duct (52) to the end piece (53) and a second channel (56) which is distinct from the first channel (55) and which connects the duct (52) to the end piece (54).

Abstract: The invention relates to a pre-vaporizing pipe (5) for a combustion chamber of a turbomachine, comprising a generally elongate main body (50), comprising a first end portion (51) delimiting an inlet duct (52) in which a fuel injector is to be mounted, the body (50) having an external surface (500) having a first length (L1) that extends from the portion (51) to at least one first fuel outlet end piece (53), and a second length (L2) that extends the first length (L1) from the end piece (53) and that extends from the end piece (53) to a second fuel outlet end piece (54). According to the invention, inside the body (50) there is a first channel (55) connecting the duct (52) to the end piece (53) and a second channel (56) which is distinct from the first channel (55) and which connects the duct (52) to the end piece (54).

Abstract: An annular wall of a combustion chamber of a turbine engine including: a cold side and a hot side; plural dilution holes to allow circulating air of the cold side to enter the hot side for dilution of an air/fuel mixture; plural cooling orifices to allow the circulating air of the cold side to enter the hot side to form a film of cooling air along the annular wall, the cooling orifices distributed spaced axially from one another and with geometric axes inclined, in an axial direction of flow of combustion gases, by an inclination angle relative to a normal to the annular wall; plural additional cooling orifices arranged directly downstream of the dilution holes and distributed spaced axially from one another, with geometric axes arranged in a plane perpendicular to the axial direction and inclined by an angle of inclination relative to a normal to the annular wall.

Abstract: An annular wall for a combustion chamber of a turbomachine. The wall presents a hot side and a cold side and includes at least one primary hole for enabling a first flow of air flowing on the cold side of the wall to penetrate to the hot side of the wall to feed combustion of fuel inside the combustion chamber, and together with a plurality of cooling holes, each having a diameter no greater than 1 mm, to enable a second flow of air flowing on the cold side of the wall to penetrate to the hot side of the wall to cool the hot side of the wall. The plurality of cooling holes can also dilute combustion gas resulting from the combustion by using the flow of air penetrating to the hot side of the wall through the cooling holes.

Abstract: A method of starting a turbine engine, including a re-try performed if a main injector has not ignited when a shaft has reached a first predetermined speed value, the re-try including: a stopping during which a starter and the ignitor device are stopped; a second ignition during which fuel is injected into the combustion chamber, the ignitor device being actuated, the second ignition being performed when a speed of rotation of the shaft reaches a second predetermined speed value; and a second starting during which the starter is actuated once more to drive the shaft in rotation.

Abstract: 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

Abstract: An annular wall for a combustion chamber of a turbomachine. The wall presents a hot side and a cold side and includes at least one primary hole for enabling a first flow of air flowing on the cold side of the wall to penetrate to the hot side of the wall to feed combustion of fuel inside the combustion chamber, and together with a plurality of cooling holes, each having a diameter no greater than 1 mm, to enable a second flow of air flowing on the cold side of the wall to penetrate to the hot side of the wall to cool the hot side of the wall. The plurality of cooling holes can also dilute combustion gas resulting from the combustion by using the flow of air penetrating to the hot side of the wall through the cooling holes.

Abstract: A method of starting a turbine engine, including a re-try performed if a main injector has not ignited when a shaft has reached a first predetermined speed value, the re-try including: a stopping during which a starter and the ignitor device are stopped; a second ignition during which fuel is injected into the combustion chamber, the ignitor device being actuated, the second ignition being performed when a speed of rotation of the shaft reaches a second predetermined speed value; and a second starting during which the starter is actuated once more to drive the shaft in rotation.

Abstract: A method for ground control of proper operation of an aeronautical turbine engine for a plane. A test includes carrying out, on the operating turbine engine and from a predetermined speed, a quick reduction in fuel flow according to a programmed decrease to evaluate flame-out resistance of the combustion chamber of the turbine engine during a quick inflight deceleration maneuver of the speed thereof.