A variable stator vane for a compressor guide vane is provided. The vane includes a blade and a pivot. The pivot includes an internal pivot element and a pivot cap. The blade and the internal pivot element are each made from a composite material. At least one contact surface of the pivot cap is metallic.

A diverter damper for controlling a gas flow in a gas duct of large cross section. The diverter damper includes a housing having an inlet and two outlets, a pivotable flap which in a first extreme position closes a first outlet and in a second extreme position closes a second outlet. The damper further includes a drive shaft connected to the pivotable flap, wherein the drive shaft extends at least partly between two opposite housing walls and through at least one of the two opposite housing walls, and at least one actuator mechanism that is located outside the housing near or against the at least one housing wall through which the drive shaft extends. The actuator mechanism includes at least one cylinder piston unit connected to the drive shaft for pivoting the flap into one of the extreme positions or into a position between the extreme positions.

A coolable wall element for a gas turbine, having a base body with a first surface subjectable to a hot gas, second surface arranged opposite of the first surface, and first seat for housing edges of an impingement plate. The wall element has an impingement plate partly inserted into the first seat located at a distance and adjacent to the second surface. A coolable wall element with extended life time is provided with the impingement plate which is removably attached to the base body having a snap-in connection with a bendable retention tab extending from the rest of the impingement plate to a free end of the retention tab, wherein the base body has a second seat for the free end of said tab, the second seat blocks the moving of the impingement plate relative to the main body when the bendable retention tab is released.

A gas turbine engine rotor assembly has a plurality of blades spaced apart from each other for rotation about an axis. Each of the blades includes a platform having an inner surface and an outer surface. The inner surfaces of adjacent platforms define a pocket having a radially outer wall, a pressure side wall, and a suction side wall. The pocket includes a leading edge wall portion and a trailing edge wall portion, and a shelf extending in a tangential direction relative to the axis from the pressure side of the pocket. The shelf is spaced apart from the radially outer wall.

A fan exit guide vane assembly for a gas turbine engine includes a plurality of guide vanes having a pressure side and a suction side extending between a leading edge and a trailing edge. The vanes further include a span extending between a root and tip with a stagger angle defined as an angle between a longitudinal axis parallel to an engine axis of rotation and a line connecting the leading edge and the trailing edge that is less than about 15.

A fuel injector for a gas turbine engine may include a center body disposed about a longitudinal axis, and a premix barrel positioned radially outwardly from the center body to define an annular passageway between the center body and the premix barrel. The annular passageway may extend from an upstream end that is configured to be fluidly coupled to a compressor to a downstream end that is configured to be fluidly coupled to a combustor. The premix barrel may include a first portion at the upstream end and a second portion at the downstream end. The first portion may include a stainless steel material, and the second portion may include a nickel based superalloy material. The second portion may be coupled to the first portion by a laser clad coupling.

A guide mechanism for a gas turbine has a casing supporting on its interior a duct delimited by radially inner and outer duct segments. A variable guide vane located at least partially within the duct can rotate about an axis of rotation to guide flow of a fluid in the duct. The radially inner duct segment is fastened to the guide vane and is retained on the casing element via the guide vane, which is mounted on the casing element via a bushing inserted in the radial direction from outside to inside into a passage opening of the casing element.

The invention relates to an instrumented airfoil for mounting in a flow passage through a turbine, compressor, or module of a turbine engine. An instrumentation member is attached to an airfoil and includes information-obtaining means received in a cutout formed in the airfoil and projecting relative to the leading edge or the trailing edge. According to the invention, the instrumentation member also comprises a holder portion that is united with the information-obtaining means and that is attached to the cutout in such a manner that its surface is in alignment respectively with the suction side wall, the pressure side wall, and the leading edge or the trailing edge of the airfoil, the transitions between these respective surfaces not presenting any setbacks.

A noise attenuation panel for a bleed flow is presented that causes a total pressure loss of the bleed flow before it is exhausted. The total pressure loss results from at least two regions in which the flow area contracts and then rapidly expands, with the rapid expansion causing mixing and turbulence rather than full total pressure recovery. This reduced pressure means that when the flow is exhausted into a flow (which may be the bypass flow of a gas turbine engine), its energy, and thus its noise, are reduced.

A method of lining a turbofan casing (30) with an array (32) of liner modules (34), including the steps of: providing a liner module (34); and providing a coupling element (48) on the casing for interlocking engagement with the liner module to maintain the liner module in position relative to the casing; and interlockingly engaging the liner module (34) and coupling element (48) by progressive movement of the liner module relative to the coupling element.