A method for manufacturing a blade or vane assembly having at least one hollow airfoil (10) for a gas turbine, profile sections (A, B, C) of this airfoil being configured (S20) on the basis of a predetermined desired torsion of the airfoil, and the airfoil being manufactured (S30) on the basis of the configured profile sections using an additive manufacturing process (S30).

A method for producing a rotating turbine blade and tuning the natural frequencies of the blade is disclosed that changes the second effective beam length of the blade thereby changing the separation between the first natural frequency and the second natural frequency of the blade.

A compressor rotor, such as a fan, for a gas turbine engine is described which includes alternating at least first and second blade types. The leading edge of the second blade types includes a leading edge tip cutback extending to the blade tip thereof. The leading edge tip cutback of the second blade type defines a chord length at the blade tip of the second blade types that is less than that of the first blades types. The first and second blade types generate different shock patterns when the fan or compressor rotor operates in supersonic flow regimes.

A bladed rotor system for a turbomachine includes a circumferential row of blades mounted on a rotor disc, and includes a plurality of under-platform dampers. Each damper is located between adjacent blade platforms. The plurality of dampers includes a first set of dampers and a second set of dampers. The dampers of the first set are distinguished from the dampers of the second set by a cross-sectional material distribution in the damper that is unique to the respective set. Dampers of the first set and the second set are positioned alternately in a periodic fashion in a circumferential direction, to provide a frequency mistuning to stabilize flutter of the blades.

A blade cascade of a turbomachine having at least one shape variation of a blade situated on the blade side in the proximity of a side wall and extending downstream, and at least one side wall contouring of the side wall or at least one second shape variation of an adjacent blade near the side wall, as well as a turbomachine, are disclosed.

An inlet guide assembly for a turbine receiving a pulsed flow, including a duct having an internal volume, and an inlet port, first outlet nozzle and second outlet nozzle each communicating with the internal volume. The inlet port is configured to receive at least part of the pulsed flow. The first and second outlet nozzles each define a respective nozzle area communicating between the internal volume and a flow path of the turbine. The first and second outlet nozzles are spaced from one another with the first outlet nozzle located closer to the inlet port than the second outlet nozzle relative to a flow direction through the duct, the nozzle area of the first outlet nozzle being smaller than the nozzle area of the second outlet nozzle. A compound engine assembly and method of introducing a pulsed flow into a flow path of a turbine are also discussed.

The invention relates to a method for producing a mistuning component. The method comprises the following steps: a) producing a container (34) having at least one chamber (36); b) producing a lid (32, 32′); c) inserting at least one impulse element into the chamber (36); d) joining the lid (32, 32′) and the container (36), wherein joining is carried out by soldering/brazing.

A circular row of non-uniformally spaced vanes includes only one first group and only one second group of adjacent vanes, unequal first and second spacing between adjacent vanes in the first and second groups, and first spacing greater than second spacing. An embodiment with second group including only three adjacent vanes. Second spacing may be about 25%-35% smaller than a nominal uniform spacing used as a design parameter for designing spacing of the non-uniformly spaced stator vanes. Circular row may be sectored. A gas turbine engine section may include one or more rings or circular rows of fixed and/or variable non-uniformally spaced vanes. Method for designing non-uniform vane spacing for circular row includes determining nominal uniform spacing and forming first spacing and second spacing from nominal uniform spacing of vanes.

A blade cascade for a turbomachine, having a number of blades (11, . . . 14; 21, . . . 25; 31, . . . 37) which include a monocrystalline material, each blade having a crystal orientation value (|α|), which is dependent on a crystal orientation of the monocrystalline material of the blade; the crystal orientation values of first blades (11, . . . 14) being less than a first limiting value and the crystal orientation values of second blades (21, . . . 25; 31, . . . 37) being at least equal to the first limiting value; and the blade cascade having at least one first sector (1), which includes at least three successive first blades (14, 12, 11, 13), and having at least one second sector (2+3; 2′+3′; 2″+3″), which includes at least three successive second blades (22, 21, 23, 24, 25; 31, 34, 36, 37, 33, 32, 35).

A turbine engine has at least two successive annular rows of stationary vanes, e.g. formed by the vanes of a nozzle and by an annular row of casing arms arranged downstream from the nozzle Each casing arm extends substantially in a radial plane passing between the trailing edges of two adjacent stationary vanes of the nozzle, and the pitch between these two stationary vanes is greater than the pitch between the other stationary vanes of the nozzle, in such a manner that the wakes formed at the trailing edges of these two stationary vanes pass respectively on either side of the corresponding casing arm.