Abstract: A blade for a turbine engine includes an aerodynamic airfoil extending radially outwards, a blade tip, a blade root, and a platform connected to the root by an upper end. A maximum radial distance between the blade tip and the platform defines a maximum extent of the bearing surface of the airfoil. The blade is partially covered by a protective strip to combat oxidation and corrosion. The strip extends between a lower limit situated on at least one portion of the platform and an upper limit situated on the blade root. The blade tip is not covered by the protective strip.

Abstract: Device for draining liquids for an aircraft engine, comprising a collector designed to collect liquids drained from the engine, said device comprising means for pumping the liquids held in the collector and for discharging said liquids, and monitoring means designed to indicate when the liquids have been collected by the collector in an anomalous manner, said monitoring means being designed to be activated when the flow rate of collected liquids is greater than the delivery of the pumping means.

Abstract: The invention relates to a turbine engine combustion assembly (20), which includes: an annular flame tube (21) including a front wall (23), a rear wall (24) and a bottom (22) arranged facing an engine shaft (30); an injection wheel (41) rotated by said engine shaft (30), partially projecting into the bottom (22) of the flame tube (21) and configured such as to spray fuel into the flame tube by centrifugation; and at least one injector (35), capable of depositing a film of fuel on said injection wheel (41), said combustion assembly being characterised in that said injector (35) is arranged so as to pass through said upstream area of the front wall (23) or the rear wall (24) of the flame tube (21), so that the injection opening thereof (37) opens into said tube (21), opposite the portion (43) of said injection wheel (41) which is located inside said flame tube (21).

Abstract: A combustion chamber, for example for a turbine engine, presenting an inner annular wall, an outer annular wall, and a pierced annular chamber end wall extending around an axis, the chamber end wall including at least one opening receiving a fuel injector, the opening being substantially centered on a circular line defining a first chamber end wall portion extending radially between the circular line and the inner annular wall, and a second chamber end wall portion extending radially between the circular line and the outer annular wall. First and second channels are inclined relative to a normal vector normal to the chamber end wall while extending tangentially, the first channels arranged to provide a flow of air in a first rotary direction, the second channels arranged to provide a flow of air in a second rotary direction opposite to the first rotary direction.

Abstract: Device (10) for draining liquids for an aircraft engine (1), comprising a collector (11) designed to collect liquids drained from the engine, characterised in that said device comprises means (13, 19) for pumping the liquids held in the collector and for discharging said liquids, and monitoring means (14) designed to indicate when the liquids have been collected by the collector in an anomalous manner, said monitoring means being designed to be activated when the flow rate of collected liquids is greater than the delivery of the pumping means.

Abstract: A low-cost combustion chamber injector device that has sturdier construction and is safer for use. Parts are configured specifically to be nested, to dispense with welding between the cellular parts. The device for injecting a fuel-air mix includes, centered on a single axis, a swirl chamber including at least two annular air suction parts, a centering guide, and a retainer ring. The annular parts have cellular conduits connecting to axial inner and outer Venturi tubes. The parts are coaxially mounted, the inner part being configured to be self-centered in the outer part by engagement of the axial and radial walls and then, after assembly, the outer part is welded to the retainer ring and to the inner part in a single radial plane.

Abstract: A combustion chamber, for example for a turbine engine, presenting an inner annular wall, an outer annular wall, and a pierced annular chamber end wall extending around an axis, the chamber end wall including at least one opening receiving a fuel injector, the opening being substantially centered on a circular line defining a first chamber end wall portion extending radially between the circular line and the inner annular wall, and a second chamber end wall portion extending radially between the circular line and the outer annular wall. First and second channels are inclined relative to a normal vector normal to the chamber end wall while extending tangentially, the first channels arranged to provide a flow of air in a first rotary direction, the second channels arranged to provide a flow of air in a second rotary direction opposite to the first rotary direction.

Abstract: A low-cost combustion chamber injector device that has sturdier construction and is safer for use. Parts are configured specifically to be nested, to dispense with welding between the cellular parts. The device for injecting a fuel-air mix includes, centered on a single axis, a swirl chamber including at least two annular air suction parts, a centering guide, and a retainer ring. The annular parts have cellular conduits connecting to axial inner and outer Venturi tubes. The parts are coaxially mounted, the inner part being configured to be self-centered in the outer part by engagement of the axial and radial walls and then, after assembly, the outer part is welded to the retainer ring and to the inner part in a single radial plane.