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).

Combustor assemblies and methods for assembling combustor assemblies are provided. For example, a combustor assembly comprises an annular inner liner and an annular outer linear, each extending generally along an axial direction. The outer liner includes an outer flange extending forward from its upstream end. The combustor assembly also comprises a combustor dome extending between an inner liner upstream end and the outer liner upstream end and including an inner flange extending forward from a radially outermost end of the combustor dome. The inner liner, outer liner, and combustor dome define a combustion chamber therebetween, and the combustor dome and a portion of the outer liner together define an annular cavity of the combustion chamber. The inner and outer flanges define an airflow opening therebetween, and a chute member is positioned within the airflow opening to define an air chute for providing a flow of air to the annular cavity.

A combustor liner panel attachment assembly includes a liner extending from a first end to a second end, and circumferentially to partially define a combustion zone. The assembly also includes a spring element located adjacent to a portion of the liner and operatively coupled to a stationary structure, the spring element having a recessed segment. The assembly further includes a protrusion feature extending radially outwardly from the liner, the protrusion feature disposed within the recessed segment of the spring element to axially retain the liner.

A turboengine as disclosed includes an outer wall structure and an inner wall structure, wherein the inner wall structure is provided at a radially inner position with respect to the outer wall structure, and each of the wall structures has a surface, the surfaces being arranged facing each other in the radial direction. At least one guide vane member includes at least one airfoil, a radially inner end and a radially outer end. The inner wall structure and the outer wall structure are jointly provided as a vane carrier unit, wherein the inner wall structure and the outer wall structure are fixedly connected to each other by at least one bridging member extending between the inner wall structure and the outer wall structure.

Fuel injection assemblies and combustors are provided. A fuel injection assembly includes a fuel injector having a first annular wall that about an axial centerline and extends from a primary inlet disposed at a first end to a primary outlet disposed at a second end. A second annular wall surrounds the first annular wall. A fuel plenum is defined between the first annular wall and the second annular wall. A fuel duct extends from a fuel outlet defined in the second annular wall to a fuel inlet. wherein the fuel duct is in fluid communication with the fuel plenum. The fuel duct includes a polygonal segment and a cylindrical inlet segment. The polygonal segment extends from the fuel outlet to the cylindrical inlet segment.

A combustor assembly includes a first combustion chamber member and a second combustion chamber member. The first and second combustion chamber members are each formed of a CMC material and include adjacently positioned coupling flanges. A mounting assembly includes a pair of mounting flanges positioned on opposite sides of the coupling flanges of first and second combustion chamber members. An attachment member of the mounting assembly clamps the pair of mounting flanges together to mount the first and second combustion chamber members.

A combustion chamber for a turbomachine having a bottom wall, at least one mixing bowl for promoting the mixing of air and fuel, mounted in an opening in the bottom wall, at least one annular baffle mounted axially downstream of the bottom wall, with respect to the direction of the gas flow within the combustion chamber, around the opening. The baffle is produced in one piece with the mixing bowl so as to form a one-piece assembly that has at least one channel for the flow of cooling air. The channel has an air inlet located upstream of the bottom wall and an air outlet located downstream of the bottom wall. The air outlet is located radially opposite the baffle.

A gas turbine engine has a movable interface that permits rotation of the inlet portion and cone. The movable interface moves within a surface of the flow sleeve that is designed to receive the movable interface. A sealing ring and groove located at the connection between the sealing ring and groove allow axial movement.

A method of manufacturing a component includes forming an inner wrap about a mandrel. The inner wrap has first and second walls joined by a base portion and an outer wall. A rod is arranged at each of the first and second walls. An outer wrap is formed about the inner wrap and the rods to form a body. Features are formed in the first and second walls.