An intermediate duct (10) for disposition between an outlet of a low-pressure compressor and an inlet of a high-pressure compressor of a turbomachine, in particular of an aircraft engine is provided, the intermediate duct including an outer wall (2) and an inner wall (3) between which are disposed an optional exit stator ring (4) and at least one strut (12) extending radially with respect to a central axis of the intermediate duct (10). The intermediate duct (10) is at least partially manufactured by additive manufacturing. A method for manufacturing such an intermediate duct (10), and a turbomachine having such an intermediate duct (10) are also provided.

A guide blade arrangement (20) for a turbomachine (1), including a guide blade airfoil (22) and a platform (21). The guide blade airfoil (22) is situated at a side (21.1) of the platform (21) facing the gas channel, an opposite side (21.2) of the platform (21) facing away from the gas channel being contoured at least in one area (30.1, 30.2) with elevations (25) and depressions (26) that follow one another in the circumferential direction (23) in relation to a longitudinal axis (2) of the turbomachine (1), and the elevations (25) and depressions (26) at the side (21.2) facing away from the gas channel being set via a platform thickness (31), taken radially in each case, that is variable in the circumferential direction (23) and that repeatedly increases and decreases with a continuous profile.

A blisk 10 for a gas turbine includes a rotor blade row 12 extending around a central axis X and, axially spaced therefrom and extending coaxially therewith, at least one annular sealing fin 11. The sealing fin has a radially outer annular portion 111 that is thickened as compared to a radially more inward annular portion 113. A compressor 1 includes a rotor and a casing 30. The casing includes at least one stator vane row having at least one abradable liner. The rotor includes at least one blisk 10, whose at least one sealing fin 11 at least partly engages in the abradable liner. A turbine is constructed analogously. A method for manufacturing a blisk 10 for a gas turbine includes producing a blisk 10 having least one annular sealing fin 11, as well as applying a coating 116 to a radially outer surface 115 of a thickened annular portion 111 of sealing fin 11.

A method of developing a suggested blend repair to an airfoil includes the steps of: (a) storing history with regard to a particular airfoil in a particular engine; (b) taking information with regard to new damage to the particular airfoil; (c) reaching an initial blend recommendation based upon step (b); (d) assessing whether the initial blend recommendation of step (c) would be appropriate to repair the new damage based upon a consideration of steps (a)-(c); and (e) reporting a final blend recommendation. An airfoil repair recommendation system is also disclosed.

Methods for repairing an airfoil having a damaged region are provided. The method can include removing the damaged portion from the airfoil to form an intermediate component. The damaged portion generally includes an original film hole having an original cross-sectional geometry. Using additive manufacturing, a replacement portion is then applied on the intermediate component to form a repaired component with the replacement portion including a rebuilt film hole having a rebuilt cross-sectional geometry that is different than the original cross-sectional geometry.

A rotary apparatus for inputting thermal energy into fluidic medium is provided, the apparatus comprises: a casing with at least one inlet and at least one outlet; a rotor comprising at least one row of rotor blades configured as impulse impeller blades arranged over a circumference of a rotor hub mounted onto a rotor shaft; at least one row of stationary nozzle guide vanes arranged upstream of the at least one row of the rotor blades, respectively; and at least one row of stationary diffuser vanes arranged downstream of the at least one row of the rotor blades, respectively. The apparatus is configured to impart an amount of thermal energy to a stream of fluidic medium directed along a flow path formed inside the casing between the inlet and the outlet by virtue of a series of energy transformations occurring when said stream of fluidic medium successively passes through the blade/vane rows formed by the nozzle guide vanes, the rotor blades and the diffuser vanes, respectively, wherein, in said apparatus, a space formed between an exit from the at least one row of diffuser vanes and an entrance to the at least one row of nozzle guide vanes in a direction of the flow path formed inside the casing between the inlet and the outlet is made variable to regulate the amount of thermal energy input to the stream of fluidic medium propagating through the apparatus. Related uses and a method for inputting thermal energy into a fluidic medium are further provided.

A vaneaxial blower system having a stator and a rotor is disclosed. Each of the stator and rotor have a hub and blades, each of the blades have a leading edge and a trailing edge, an angle theta defined by a line extending from the respective leading edges to the respective trailing edges and the respective hub and an angle beta defined as an angle measured between a camber line between a first blade and a second blade and a horizontal tangential line that extends through respective leading edges of the first and second blade.

Provided is a vibration reduction device for stator vanes (30) positioned behind rotor blades (28) in a turbo machine, comprising an annular base member (80) having a cylindrical shape concentric around a central axis of the casing and supporting base ends of the stator vanes which extend radially inward from an inner circumferential surface of the base member, an elastomeric damping member (100) surrounding and in slidable contact with an outer circumferential surface of the base member, and a preloading member (102) surrounding an outer circumferential surface of the elastomeric damping member and configured to apply a preload directed radially inward to the elastomeric damping member.

The invention relates to part of a turbine engine comprising two arms passing through a stream of the turbine engine, wherein each arm comprises an outer surface and an aerodynamic linking device. The aerodynamic linking device comprises fairings extending between the two arms, compressible interface means interposed between the fairings and means for retaining the fairings in place by pressure in relation to the arms, which compress the interface means.

Provided is a radial diffuser that includes a housing, a plurality of diffuser vanes, and a plurality of deswirl vanes and at least one vane extension providing a service routing. Each of the vane extensions is disposed after a radial section and may extend into or through a transition and into the deswirl cascade. At least a portion of the vane extensions include a service passage extending therethrough. Each service passage is configured to allow a service conduit to extend therethrough without adversely crossing either a diffusion flow passage or a transition flow passage.