Abstract: The invention relates to a combustion chamber for a turbomachine. The combustion chamber comprises at least one spark plug, at least one fuel injector of a first type of injector that is located facing the spark plug, and at least one fuel injector of a second type. Each fuel injector comprises a mobile sealing member that is configured to open when the fuel pressure exceeds an opening threshold. The opening threshold of the mobile sealing member of the injector of the first type is strictly less than the opening threshold of the mobile sealing member of the injector of the second type.

Abstract: A combustion chamber of a turbomachine includes an annular outer casing and a flame tube connected to the outer casing. The flame tube includes an annular inner wall and an annular outer wall defining a first, radial portion at the inlet of the flame tube and a second, axial portion at the outlet of the flame tube. The first portion extends towards the second portion forming an elbow between the inlet and the outlet of the flame tube.

Abstract: The invention relates to a combustion chamber for a turbomachine, such as an aircraft turbojet engine or turboprop engine, comprising an internal annular shroud and an external annular shroud which at their upstream ends are connected by an annular chamber end wall (48), said chamber comprising deflectors (50) mounted upstream of the annular chamber end wall (48). Injectors (19) are mounted in sleeves (44) at least one of which comprises a radially annular flange (66) which is designed to slide radially between the chamber end wall (48) and the deflector (50) and which is blocked axially between the chamber end wall (48) and the deflector (50).

Abstract: The invention relates to a combustion chamber for a turbomachine. The combustion chamber comprises at least one spark plug, at least one fuel injector of a first type of injector that is located facing the spark plug, and at least one fuel injector of a second type. Each fuel injector comprises a mobile sealing member that is configured to open when the fuel pressure exceeds an opening threshold. The opening threshold of the mobile sealing member of the injector of the first type is strictly less than the opening threshold of the mobile sealing member of the injector of the second type.

Abstract: A combustion chamber for a turbine engine, including an annular end wall provided with injection systems each centered on a respective axis and each having an upstream end forming a bushing for receiving a head of a fuel injector, and an annular shroud covering the end wall and including injector ports respectively arranged facing the injection systems, wherein the annular shroud includes air intake ports separate from the injector ports, and the bushing of each injection system crosses the corresponding injector port and includes at its upstream end a collar having a free end remote from the axis of the injection system by a first distance greater than a second distance separating a rim of the injector port from the axis.

Abstract: A combustion chamber of a turbomachine includes an annular outer casing and a flame tube connected to the outer casing. The flame tube includes an annular inner wall and an annular outer wall defining a first, radial portion at the inlet of the flame tube and a second, axial portion at the outlet of the flame tube. The first portion extends towards the second portion forming an elbow between the inlet and the outlet of the flame tube.

Abstract: The invention relates to a combustion chamber for a turbomachine, such as an aircraft turbojet engine or turboprop engine, comprising an internal annular shroud and an external annular shroud which at their upstream ends are connected by an annular chamber end wall (48), said chamber comprising deflectors (50) mounted upstream of the annular chamber end wall (48). Injectors (19) are mounted in sleeves (44) at least one of which comprises a radially annular flange (66) which is designed to slide radially between the chamber end wall (48) and the deflector (50) and which is blocked axially between the chamber end wall (48) and the deflector (50).

Abstract: A combustion chamber for a turbine engine, including an annular end wall provided with injection systems each centred on a respective axis and each having an upstream end forming a bushing for receiving a head of a fuel injector, and an annular shroud covering the end wall and including injector ports respectively arranged facing the injection systems, wherein the annular shroud includes air intake ports separate from the injector ports, and the bushing of each injection system crosses the corresponding injector port and includes at its upstream end a collar having a free end remote from the axis of the injection system by a first distance greater than a second distance separating a rim of the injector port from the axis.

Abstract: A fuel injection device is provided. The device includes a pilot injector opening out into a venturi and an annular row of multipoint injection orifices opening out axially in a downstream direction and radially outside the venturi, the venturi being connected at its downstream end to a separation wall. The outer peripheral edge of the separation wall has a diameter greater than the diameter of the annular row of injection orifices, whereby the separation wall forms a thermal protection screen for these orifices against heat radiation coming from the chamber.

Abstract: An air and fuel injection system for the back of a turbomachine's annular combustion chamber, including a central injector and a peripheral annular fuel injector including at least one fuel injection aperture made in an annular wall and emerging in a peripheral annular channel separated from a central channel by the annular wall and having an annular space of admission of air. The system also includes multiple air ejection apertures made in the annular wall downstream from the fuel ejection aperture or apertures in reference to the flow of the air stream, for an additional injection of air into the channel.

Abstract: A combustion chamber for an aviation turbine engine, the combustion chamber being annular about a longitudinal axis and including an outer side wall, an inner side wall, and an annular chamber end wall connecting one end of the outer side wall to one end of the inner side wall. The outer side wall includes, distributed along its circumference, spark plugs, primary holes, and dilution holes situated downstream from the primary holes in the direction of the longitudinal axis. The primary holes situated in each of the adjacent zones adjacent to one of the spark plugs present a configuration that is different from the configuration of the primary holes situated outside the zones, such that the supply of air in the adjacent zones is different from the supply of air outside the zones.

Abstract: Annular combustion chamber (10) to be fitted on a turbomachine and comprising a chamber end wall (22), a plurality of air and fuel injection systems (32) circumferentially distributed around an axis (34) of the combustion chamber (10) and mounted on said chamber end wall (22), and, an air manifold (100) associated with each injection system (32), comprising at least one wall (96, 98) mounted on the chamber end wall (22) and projecting in the upstream direction to form an obstacle to a circumferential airflow around the axis (34) of the combustion chamber (10), and an air inlet opening (88) formed at the upstream end of the air manifold (100) and opening radially outwards from an axis (44) of said injection system.

Abstract: A fuel injector device for an annular combustion chamber including a pilot circuit feeding an injector and a multipoint circuit feeding injection orifices formed in a front face of an annular chamber, an annular ring being mounted in the annular chamber to define therein a fuel feed circuit for the injection orifices and a cooling circuit operating by passing the fuel feeding the injector and extending over the front face of the chamber in immediate vicinity of the injection orifices.

Abstract: A fuel injector device for an annular combustion chamber of a turbine engine, including a pilot circuit feeding an injector and a multipoint circuit feeding injection orifices formed in a front face of an annular ring mounted in an annular chamber, together with a thermal insulation mechanism insulating the front face and including an annular cavity formed around the injection orifices between the front face of the annular ring and a front wall of the annular chamber and configured to be filled in operation with air or with coked fuel.

Abstract: A multipoint injector for a turbomachine according to which any risk of fuel coking is eliminated is disclosed. The multipoint fuel which is liable to stagnate inside the circuit thereof is cooled uniformly, due to the formation of continuous baffles which each communicate with at least one separate circulation channel and of which the peripheral baffles open out into a fuel admission chamber arranged in a zone diametrically opposing the circulation channels and which communicates with the injection nozzle for pilot fuel in order to achieve uniform supply and cooling of the injector.

Abstract: A fuel injection device (110) for a turbine engine, the device comprising a pilot injector opening out into a venturi (120) and an annular row of multipoint injection orifices (136) opening out axially in a downstream direction and radially outside the venturi, the venturi being connected at its downstream end to a separation wall (140), the device being characterized in that the outer peripheral edge (142) of the separation wall has a diameter (B?) greater than the diameter of the annular row (A) of injection orifices, whereby the separation wall forms a thermal protection screen for these orifices against heat radiation coming from the chamber.

Abstract: A fuel injector device for an annular combustion chamber including a pilot circuit feeding an injector and a multipoint circuit feeding injection orifices formed in a front face of an annular chamber, an annular ring being mounted in the annular chamber to define therein a fuel feed circuit for the injection orifices and a cooling circuit operating by passing the fuel feeding the injector and extending over the front face of the chamber in immediate vicinity of the injection orifices.

Abstract: A fuel injector device for an annular combustion chamber of a turbine engine, including a pilot circuit feeding an injector and a multipoint circuit feeding injection orifices formed in a front face of an annular ring mounted in an annular chamber, together with a thermal insulation mechanism insulating the front face and including an annular cavity formed around the injection orifices between the front face of the annular ring and a front wall of the annular chamber and configured to be filled in operation with air or with coked fuel.

Abstract: A combustion chamber for an aviation turbine engine, the combustion chamber being annular about a longitudinal axis and including an outer side wall, an inner side wall, and an annular chamber end wall connecting one end of the outer side wall to one end of the inner side wall. The outer side wall includes, distributed along its circumference, spark plugs, primary holes, and dilution holes situated downstream from the primary holes in the direction of the longitudinal axis. The primary holes situated in each of the adjacent zones adjacent to one of the spark plugs present a configuration that is different from the configuration of the primary holes situated outside the zones, such that the supply of air in the adjacent zones is different from the supply of air outside the zones.

Abstract: A multipoint injector for a turbomachine according to which any risk of fuel coking is eliminated is disclosed. The multipoint fuel which is liable to stagnate inside the circuit thereof is cooled uniformly, due to the formation of continuous baffles which each communicate with at least one separate circulation channel and of which the peripheral baffles open out into a fuel admission chamber arranged in a zone diametrically opposing the circulation channels and which communicates with the injection nozzle for pilot fuel in order to achieve uniform supply and cooling of the injector.