A tandem rotor disk apparatus may include a rotor disk body concentric about an axis. The tandem rotor disk apparatus may also include a first blade extending radially outward of the rotor disk body and a second blade extending radially outward of the rotor disk body. The first blade may be offset from the second blade in a direction parallel to the axis. The tandem rotor disk apparatus may be implemented in a gas turbine engine with no intervening stator vane stages disposed between the first blade and the second blade. The tandem rotor disk apparatus may include two separate rotor disk bodies.

A bladed assembly for a turbine of a turbomachine comprises an outer platform comprising, at an axial end, a transverse end surface and a sealing device protruding from the transverse end surface and comprising at least two sealing ribs (102) extending transversely to an axis of the bladed assembly and each having two respective sides which, in an axial cross-section, are radially and axially inclined in the same direction with respect to the axis. The radial direction of the inclination of the sides of the sealing ribs, in an axial cross-section, is the same for at least two of the sealing ribs. This particularity confers flexibility on each of the sealing ribs, which allows to favour a surface contact between the sealing ribs and a radial stator wall facing which the transverse end surface, and in particular the sealing device, are intended to be arranged.

A turbine nozzle includes an airfoil shape. The airfoil shape may have a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The airfoil profile sections at Z distances may be joined smoothly with one another to form a complete airfoil shape.

Provided is a steam turbine in which, among a plurality of stages thereof, a stage that is disposed furthest upstream is a speed-adjusting stage, at least one stage that is disposed downstream of the speed-adjusting stage is a medium-pressure stage, and at least one stage that is disposed downstream of the medium-pressure stage is a low-pressure stage. The speed-adjusting stage is an impulse stage. The medium-pressure stage is a medium reaction degree impulse stage in which the degree of reaction is a medium degree of reaction of 10-40%. The low-pressure stage is a reaction stage having a higher degree of reaction than the medium-pressure stage.

A gas turbine engine includes a compressor and a turbine; a plurality of rotor blades positioned in the compressor or the turbine and arranged in a plurality of stages along an axial direction; an outer drum positioned at least partially radially outward of the plurality of rotor blades and rotatable with the plurality of rotor blades; and a casing extending over the outer drum and defining a plurality of air cavities with the outer drum, the plurality of air cavities arranged in series airflow communication along the axial direction.

A fan for a gas turbine engine is provided. The fan includes a rotor including at least one rotor stage having a rotatable disk defining a flowpath surface and an array of rotor airfoils extending outward from the flowpath surface; an array of splitter airfoils extending outward from the flowpath surface, wherein the splitter airfoils and the rotor airfoils are disposed in a sequential arrangement; and a shroud extending between the rotor airfoils and the splitter airfoils.

A tandem rotor disk apparatus may include a rotor disk body concentric about an axis. The tandem rotor disk apparatus may also include a first blade extending radially outward of the rotor disk body and a second blade extending radially outward of the rotor disk body. The first blade may be offset from the second blade in a direction parallel to the axis. The tandem rotor disk apparatus may be implemented in a gas turbine engine with no intervening stator vane stages disposed between the first blade and the second blade. The tandem rotor disk apparatus may include two separate rotor disk bodies.

A system for increasing the efficiency of low pressure axial fans, includes, a low pressure axial fan including multiple first blades, and, a vortex generator operably coupled to the low pressure axial fan via a shaft. In use, the vortex generator includes multiple second blades. In further use, the low pressure axial fan is configured to operate in an operational mode and the vortex generator is configured to operate in a stationary mode.

Airfoil assemblies for gas turbine engines are provided. For example, an airfoil assembly comprises an airfoil, an inner band defining an inner opening shaped complementary to an inner end of the airfoil, and an outer band defining an outer opening shaped complementary to an outer end of the airfoil. The airfoil inner end is received with the inner opening, and the airfoil outer end is received within the outer opening. A strut extends radially through an airfoil cavity. A first pad is defined at a first radial location within the cavity. A second pad is defined within the cavity at a second, different radial location. In some embodiments, the airfoil assembly inner band includes a first inner flange, through which the inner band is secured to a support structure, and the outer band includes a first outer flange, through which the outer band is secured to a support structure.

The present application provides a turbine blade. The turbine blade includes a root section with a first curved section, a tip section with a second curved section, and number of mean sections positioned between the root section and the tip section. The mean sections each include a substantially prismatic shape.