Turbine assemblies for gas turbine engines are provided. In one embodiment, a turbine assembly includes a row of circumferentially adjacent turbine rotor blades, each blade in the row of blades including an airfoil defining a chord extending axially between opposite leading and trailing edges of the airfoil. The chords in a mid-span region of each blade are shorter than the chords in span regions adjacent to the blade root and the blade tip. In another exemplary embodiment, a turbine assembly includes a row of circumferentially adjacent turbine stator vanes, each vane in the row of vanes including an airfoil defining a chord extending axially between opposite leading and trailing edges of the airfoil. The chords in a mid-span region of each vane are shorter than the chords in span regions adjacent to the vane root and the vane tip. A turbine rotor blade for a gas turbine engine also is provided.

A turbine engine assembly including a housing and a bladed wheel rotatable within the housing. The bladed wheel includes at least one blade including a head opposite the housing. The head includes a magnet and the housing includes a first and second electrical conductor. Each electrical conductor is configured to generate, across terminals thereof, an electrical voltage induced by the magnet of the head opposite the housing and that represents vibrations to which the head of the blade is subjected when the bladed wheel is rotated. The first electrical conductor includes a first central portion extending around the rotational axis of the bladed wheel and includes two mutually facing ends, and the second electrical conductor includes a second central portion passing through a space left by the first central portion between the two ends thereof.

An apparatus and method of cooling a hot portion of a gas turbine engine, such as a rotor disk, by having a vane assembly with a cooling air passage and a flow control insert located within the cooling air passage defining a conduit. A turning nozzle is mounted to the vane and has a turning passage with an inlet and an outlet, the turning nozzle is fluidly coupled to the flow control insert outlet.

A turbomachine blade of the type having a metal lower coupling root, a metal upper coupling head, and a metal airfoil-shaped oblong member designed to connect the coupling root rigidly to the coupling head; the airfoil-shaped oblong member having a substantially airfoil-shaped main plate-like element connected to the coupling root and to the coupling head, and which is divided into: a lower connecting fin cantilevered from and formed in one piece with the coupling root; an upper connecting fin cantilevered from and formed in one piece with the coupling head; and a center plate-like body, which is located between the lower and upper connecting fins, is shaped/designed to form an extension of the lower and upper connecting fins, and is butt-welded to, to form one piece with, the lower and upper connecting fins.

A method for repairing an airfoil of an axial turbomachine in which material is deposited onto the airfoil by means of deposition welding, the airfoil being cooled during the deposition welding, is provided. A cooling collar including at least one cooling channel which has a coolant inlet and a coolant outlet and through which a coolant flows in the intended state, is also provided. The cooling collar also includes multiple cooling elements which are arranged along an inner circumference of the cooling collar and adjacently to the at least one cooling channel, the cooling elements resting against an object to be cooled, in particular an airfoil to be cooled, in the intended state.

A plurality of slots with different widths are provided in a plurality of line on a stationary blade surface. More specifically, the steam turbine stationary blade has a hollow nozzle with a penetrating space, which is connected with a diaphragm outer ring or inner ring, and a plurality of suction slots extending radially which are arranged on the blade surface. At a position where a water film deposited to the blade surface is thick, the width of a slot is smaller and at a position where the water film is thin, the width of a slot is larger.

Structures for reducing forward loads in compressors are described. A compressor includes inner and outer circumferential support structures positioned concentrically around a central axis, and an aft-most stage including a vane extending radially inward from the outer circumferential support structure. An axial length of the aft-most stage is defined by a spacer arm of the inner circumferential support structure. The vane includes a root, a tip, and a trailing edge extending between the root and the tip. A ratio of a first radial distance between a first point located at an intersection of the tip of the vane and the trailing edge to a radially inner wall of the spacer arm of the inner circumferential support structure to a second radial distance between the first point and a second point located at an intersection of the root of the vane and the trailing edge is between 0.95 and 4.5.

A turbine section of a gas turbine engine comprises a gas path having an outer boundary wall. A circumferential array of turbine blades projects radially into the gas path. Each turbine blade extends in span from a hub to a tip and in chord from a leading edge to a trailing edge. A circumferential array of tip flow vanes extends radially inward from the outer boundary wall with a span corresponding generally to a radial depth of a tip leakage flow region of the turbine blades. The tip flow vanes are disposed downstream of the circumferential array of turbine blades adjacent to the trailing edge of the turbine blades.

The invention relates to a guide vane for a gas turbine, comprising an airfoil, a platform arranged at a radial end of the airfoil, an upstream flange extending radially from the platform, and a downstream flange extending radially from the platform, wherein the flanges, together with a section of the platform lying between the flanges, bound a groove extending in the circumferential direction of the gas turbine for the arrangement of a damping element. A surface of the section of the platform bounding the groove is arched radially at least in regions thereof in the direction of an opening of the groove. The invention further relates to a guide vane cluster, a housing for a gas turbine, as well as to a gas turbine.

A stator vane is comprised of: an airfoil section elongated in a radial direction relative to an axis; an outer band section continuous to an outer end of the airfoil section and bent in a circumferential direction relative to the axis; a first hook section continuous to a leading end in the axial direction of the outer band section and bent outward in the radial direction; a second hook section continuous to a trailing end in the axial direction of the outer band section and bent outward in the radial direction; an inner band section continuous to an inner end of the airfoil section and bent in the circumferential direction; a flange section continuous to an end in the axial direction of the inner band section and bent inward in the radial direction; and a reinforcement fiber fabric continuous throughout these sections and unitized with a ceramic.