A turbine airfoil (10) includes a trailing edge coolant cavity (41f) located in an airfoil interior (11) between a pressure sidewall (14) and a suction sidewall (16). The trailing edge coolant cavity (41f) is positioned adjacent to a trailing edge (20) of the turbine airfoil (10) and is in fluid communication with a plurality of coolant exit slots (28) positioned along the trailing edge (20). At least one framing passage (70, 80) is formed at a span-wise end of the trailing edge coolant cavity (41f). The airfoil (10) further includes framing features (72A-B, 82A-B) located in the framing passage (70, 80). The framing features are configured as ribs (72A-B, 82A-B) protruding from the pressure sidewall (14) and/or the suction sidewall (16). The ribs (72A-B, 82A-B) extend partially between the pressure sidewall (14) and the suction sidewall (16).

A rotor stage (10) of a turbine engine includes a circumferential row of rotor segments (12), each including: first and second endwalls (14, 16) spaced apart radially, and a first and second sidewalls (18, 20) extending radially between the first and second endwalls (14, 16) and spaced apart circumferentially. The first and second endwalls (14, 16) and the first and second sidewalls (18, 20) define therewithin a flow passage (22) for hot gas. Circumferentially adjacent segments (12a, 12b) mate along a respective split-line (24) extending along an interface between the first sidewall (18) of a first segment (12a) and the second sidewall (20) of a second circumferentially adjacent segment (12b). A composite airfoil structure (26) is thereby defined having a pressure sidewall (18) formed by the first sidewall (18) of the segment (12a) and a suction sidewall (20) formed by the second sidewall (20) of the second segment (12b). The first and second endwalls (14, 16) are respectively configured as a platform (14) and a tip shroud (16) of the segment (12).

A ceiling fan or blade thereof can include a fan motor for rotating the blade. The blade can include an airfoil body having an outer surface extending between a leading edge and a trailing edge, and a root and a tip. The blade can be separated into three distinct cross sections including a first cross section as a lifting cross section, a second cross section as a flat cross section, and a third cross section as a transition section between the first and second cross sections.

Component for gas turbine engines are described. The components include an airfoil body having leading and trailing edges and pressure and suction sides. The airfoil has a leading edge cavity located proximate the leading edge defined between the leading edge and a separator rib and between the pressure side and the suction side. An insert member is installed within the leading edge cavity. The insert member has one or more metering flow apertures at an aft end and one or more impingement apertures at a forward end and at least one axially extending rib along an exterior surface thereof. At least one axial extending flow channel passage is defined along the axial extending rib between an exterior of the insert member and an interior of the airfoil body. Air flow through the metering flow apertures flows into the axial extending flow channel passage and flows forward toward the leading edge.

A rotor for a turbofan booster section associated with a fan section of a gas turbine engine includes a rotor blade having an airfoil having a leading edge, a trailing edge and a mean camber line. The airfoil has a delta inlet blade angle defined as a difference between a local inlet blade angle defined in a spanwise location, and a root inlet blade angle defined at the root. The delta inlet blade angle decreases in the spanwise direction from the root to a minimum value at greater than 10% span and from the minimum value, the delta inlet blade angle increases to the tip. The rotor includes a rotor disk coupled to the rotor blade configured to be coupled to the shaft or the fan to rotate with the shaft or the fan, respectively, at the same speed as the shaft or the fan.

Turbines and associated components, systems, and methods are described. In some embodiments, the turbine blades and turbines are configured to convert kinetic energy present in fluid (e.g., water) to other forms of energy (e.g., in a hydrokinetic energy system in a river or ocean) relatively efficiently and/or at relatively low cut-in speeds. The turbine blades may have a shape and/or include structural features that contribute at least in part to relatively high efficiency and/or relatively low cut-in speeds. In some embodiments, the turbine blades have a geometry similar to the geometry of a maple seed.

A platform for a gas turbine engine according to an example of the present disclosure includes, among other things, a platform body that extends in a circumferential direction between first and second sidewalls to define a gas path surface, and opposed rows of flexible retention tabs that extend from the platform body and are dimensioned to wedge the platform between adjacent airfoils.

A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (α3−α1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.

A turbine blade having a gas film cooling structure with a composite irregular groove. The turbine blade has a hollow structure, and a plurality of first grooves which are recessed grooves are provided on an outer surface thereof. A plurality of discrete holes A extending to an inner surface of the turbine blade are provided at the groove bottom of each first groove. The first groove is an irregular groove, and includes at least two portions in a depth direction. A portion having a depth H.sub.1 from the groove bottom of the first groove is a first portion, and the rest thereof is a second portion. At least one side surface of the second portion is formed by expanding in lateral direction from a corresponding side surface of the first portion.

A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.