A guide vane assembly for a stator vane stage comprises at least two aerofoil members joined together using a vane assembly attachment web. The vane assembly attachment web may be provided at the inner and/or outer radius of the aerofoil members. The vane assembly attachment web allows the guide vane assembly to be attached to an inner and/or outer attachment ring, thereby forming a stator vane stage. Such an arrangement may allow composite fiber reinforced guide vane assemblies to be readily assembled together to form stator vane stages. Stator vane stages that comprise such composite fiber reinforced guide vane assemblies may be lighter than conventional metallic stator vane stages.

The invention relates to a method for manufacturing a housing of a turbomachine, in particular a gas turbine. The method comprises at least the steps: providing a housing blank (10), manufacturing a housing element (14), producing an assembly opening (12) corresponding to the housing element (14) in the housing blank (10), arranging the housing element (14) in the assembly opening (12), and joining the housing element (14) to the housing blank (10) by means of a welding method. In addition, the invention relates to a turbomachine housing.

Walls of gas turbine engine casings, fan cases, and methods for forming walls, e.g., fan case walls, are provided. For example, a wall comprises a plurality of composite plies arranged in a ply layup. The wall is annular and circumferentially segmented into a plurality of regions that include at least one first region and at least one second region. The ply layup in the first and second regions is different such that the ply layup is non-axisymmetric. An exemplary fan case comprises an annular inner shell, a filler layer, an annular back sheet, and an annular outer layer. The back sheet is circumferentially segmented into a plurality of regions, including at least one first region and at least one second region, and comprises a plurality of composite plies arranged in a ply layup that is different in the first and second regions such that the ply layup is non-axisymmetric.

A fluid transfer assembly for use in a gas turbine engine includes a rotor shaft, a stationary assembly circumscribing the rotor shaft, and a rotating component coupled to the rotor shaft and positioned radially inward of the stationary assembly. The rotating assembly includes a hub coupled to the rotor shaft, a plurality of rotor blades coupled to the hub, and a shroud coupled to the plurality of rotor blades.

A nacelle for a turbofan engine includes a fan case, an inner barrel and a nose lip generally concentrically disposed about an axis. The inner barrel projects forward from and is engaged to the fan case at a first split line. The nose lip is engaged to the inner barrel at a second split line, and extends substantially radially outward directly from the second split line, then substantially axially rearward to the fan cowl as one unitary piece.

An airfoil assembly includes first and second fiber-reinforced composite airfoil rings. Each ring is comprised of inner and outer platform sections, a suction side wall that extends between the inner and outer platforms, and a pressure side wall that extends between the inner and outer platforms. the first ring mates at an interface with the second ring such that the suction side wall of the first ring and the pressure side wall of the second ring together form an airfoil that circumscribes an internal cavity. The interface along the airfoil includes a gap that is configured to emit cooling air from the internal cavity.

An additively manufactured solid fuel grain for a hybrid rocket engine having a cylindrical shape, defining a center combustion port and comprising a stack of fused layers of polymeric material suitable for hybrid rocket fuel. Each layer is formed as a plurality of fused abutting concentric beads of solidified material arrayed around the center port. An oxidizer is introduced into the solid fuel grain through the center port, with combustion occurring along the exposed surface area of the solid fuel grain center port wall. Each concentric bead possesses a surface pattern that increases the combustion surface area and when stacked forms a rifling pattern of undulations that induces oxidizer-fuel gas axial flow to improve combustion efficiency. The port wall surface pattern persists during the rocket engine's operation as the fuel phase changes from solid to gas and is ablated.

A disclosed fan drive gear system includes a sun gear rotatable about an axis of rotation, a plurality of intermediate gears rotatable about an intermediate gear rotation axis in meshing engagement with the sun gear and a ring gear circumscribing the intermediate gears. A bearing assembly supports at least one of the plurality of intermediate gears and includes a first beam extending in a first direction and a second beam extending from an end of the first beam in a second direction. The bearing surface supported on the second beam such that first and second beams are configured to maintain the bearing surface substantially parallel to the intermediate gear rotation axis during operation.

The present invention relates to a blower vane of a turbomachine, comprising a structure made of composite material which comprises a fibrous reinforcement and a matrix in which the fibrous reinforcement is embedded. The vane comprises a first rigidification insert which extends in a first direction and at least one second insert which is connected to the first in a second direction which is non-collinear with the first direction, the first insert projecting from the vane in order to be connected to a disc of a turbomachine element.

A turbofan engine includes fan section including a plurality of fan blades, a gear train, a low spool including a low pressure turbine and a low pressure compressor, the low pressure turbine driving the plurality of fan blades through the gear train, and a high spool including a high pressure turbine driving a high pressure compressor. A fan nacelle at least partially surrounds a core nacelle to define a fan bypass flow path. A fan variable area nozzle is in communication with the fan bypass flow path and defines a fan nozzle exit area between the fan nacelle and the core nacelle. The fan variable area nozzle varies the fan nozzle exit area.