Each fuel injector has a main nozzle and pilot nozzle. A main fuel manifold to supply fuel to the main nozzle of each fuel injector and pilot fuel manifold to supply fuel to the pilot nozzle of each fuel injector. The main fuel manifold includes plurality of flexible main fuel pipes. Also, plurality of cross-piece connectors including a stem and four arms. The stem of each connector mounted on one fuel injector. The first and third arms extend in opposite directions from the stem and second and fourth arms extend in opposite directions from the stem. Each fuel pipe interconnects the first arm of one connector with a third arm of an adjacent connector and each connector so the first and third arms are at an angle relative perpendicular to the axis of the annular combustion chamber casing so the main fuel manifold extends around the combustion chamber casing sinusoidally.

A profiled structure for an aircraft or turbomachine is elongated in a direction in which the structure has a length exposed to an airflow and includes serrations defined by successive teeth and depressions. The serrations may be transverse to a leading edge and/or a trailing edge of the profiled structure and in the direction of elongation. Along the profiled leading edge and/or profiled trailing edge, the successive teeth and depressions may extend only over a part of the length exposed to the flow. The amplitude and/or spacing of the teeth may vary monotonically except for the few teeth nearest each end of the part, with a remaining part of the length being smooth.

A fan (1) for a smoke and/or vapour extraction system that includes an impeller (3) located within a diffuser or volute (2), the impeller (3) being driven by an electric motor (24); the impeller (3) having a body (20) having a flat part (21) from which a plurality of radially arranged blades (33) extend; these blades are separated from each other by a variable distance based on a periodic function. In addition to this, the diffuser (2) has an air delivery or discharge conduit (5) having a cross-section increasing from an inlet to its outlet (6).

A damper system may include an inner damper having a first annular geometry and a textured surface. An outer damper may also have an annular geometry and be disposed about the inner damper. The outer damper may have a textured surface. The textured surfaces of the inner damper and outer damper may define a cavity that is configured to contain a viscous fluid.

A seal assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a seal that has an elongated seal body having a seal face that bounds a gas path and an opposed impingement face. The seal body defines an internal cavity extending in a circumferential direction between opposed mate faces and extending in a radial direction between walls of the seal body defining the seal and impingement faces. A baffle divides the internal cavity into at least a first region and a second region. The first region has a first section extending transversely from a second section. The first section has a component in the radial direction, and the second section has a component in an axial direction such that the second region is defined between the baffle and the walls of the seal body defining the seal face. A method of sealing is also disclosed.

An aerofoil component defines an in use leading edge and a trailing edge. The leading edge has at least one serration defining an apex and a nadir. The leading edge has a generally chordwise extending slot located at the nadir of each serration.

A turbine rotor blade is additively manufactured and includes an airfoil body with a radially extending chamber for receiving a coolant flow, a tip end at a radial outer end of the airfoil body, and a shank at a radial inner end of the airfoil body. The radially extending chamber extends at least partially into the shank to define a shank inner surface. An integral impingement cooling structure is within the radially extending chamber. The integral impingement cooling structure allows an exterior surface of a hollow body thereof to be uniformly spaced from the airfoil inner surface despite the curvature of the chamber. The turbine rotor blade has impingement cooling throughout the blade.

Aspects of the disclosure generally relate to a turbine center frame for a turbine engine through which a flow path extends. The turbine center frame can include an inner wall radially spaced from an outer wall, with the inner and outer walls extending between an inlet and an outlet, and with the outlet downstream of the inlet with respect to the flow path. A set of circumferentially-spaced airfoils can extend between the inner wall and the outer wall.

An abradable sealing element comprises a substrate and a sealing structure. The sealing structure comprises one or more wall structures extending from the substrate and defining at least one open cell which is filled with abradable material. The one or more wall structures are formed by additive-layer, powder-fed, laser-weld deposition onto the substrate. The one or more wall structures are formed from nickel-based superalloy and constitute from about 10% to about 50% of the total volume of the sealing structure.

An apparatus and method of a gas turbine engine comprising a rotor having at least one disk with a rotor defining an axial face and a stator having at least one ring with a stator axial face confronting the rotor axial face, with terminal portions of the axial faces forming a fluid outlet there between. A recess formed in one of the axial faces defines a buffer cavity into which a wing extends from the other of the axial faces and having a surface confronting the fluid outlet. A flow reverser is further provided within at least one of the surface or the terminal portion of the other of the axial faces.