A component according to an exemplary aspect of the present disclosure includes, among other things, a platform and a feed cavity that feeds a cooling fluid inside of the platform. The feed cavity includes a leg portion and a main body portion that extends from the leg portion inside of the platform. A cooling cavity is inside the platform and in fluid communication with the feed cavity.

An airfoil may comprise an airfoil body having a leading edge and a trailing edge. A heat pipe may be disposed within the airfoil. The heat pipe may include a vaporization section and a condensation section. The vaporization section may be disposed within the airfoil body and may be configured to remove heat from the trailing edge. The second cooling apparatus may be disposed within the airfoil body and may be configured to remove heat from the leading edge.

The present disclosure relates to a blade and a gas turbine, and a blade may include: a body part; an injection hole comprising a communication part connected to an internal flow passage formed inside the body part; and an extension connecting the communication part and a surface of the blade to each other and having a width gradually increasing from the communication part the surface of the blade and configured to inject a cooling fluid to the surface of the blade; and an injection guide means disposed in the extension and configured to uniformly inject the cooling fluid in an expanding manner so as to form a cooling film on the surface of the blade. According to the present disclosure, it is possible to form the cooling film wider on the surface of the blade, by preventing a problem of a separation that occurs during when the cooling fluid flows through materialization of the uniform injection of the cooling fluid to the surface of the blade.

A heat transfer apparatus for a gas turbine engine includes: a component having a wall structure defining a flow bounding surface; a chamber formed in the component, the chamber including a wicking structure, a vapor channel, and a working fluid.

The present disclosure is directed to a rotating component for a turbine engine. The rotating component defines a surface and includes a heat pipe positioned on the surface of the rotating component or within the rotating component. The heat pipe includes a working fluid and an outer perimeter wall.

Ceramic matrix composite (CMC) airfoils and methods for forming CMC airfoils are provided. In one embodiment, an airfoil is provided that includes opposite pressure and suction sides extending radially along a span and opposite leading and trailing edges extending radially along the span. The leading edge defines a forward end of the airfoil, and the trailing edge defines an aft end of the airfoil. A trailing edge portion is defined adjacent the trailing edge at the aft end, and a pocket is defined in and extends within the trailing edge portion. A heat pipe is received in the pocket. A method for forming an airfoil is provided that includes laying up a CMC material to form an airfoil preform assembly; processing the airfoil preform assembly; defining a pocket in a trailing edge portion of the airfoil; and inserting a heat pipe into the pocket.

A cooling system for cooling a fluid reaction apparatus of a gas turbine engine includes a vapor cooling subsystem and a film cooling subsystem. The vapor cooling subsystem has a vaporization section and a condenser section for cooling a portion of the fluid reaction apparatus. The condenser section is cooled by a fluid. The film cooling subsystem is configured for cooling a portion of the fluid reaction apparatus by discharging fluid out of openings defined in the fluid reaction apparatus. At least a portion of the fluid used to cool the condenser section of the vapor cooling subsystem is discharged out of the openings of the film cooling subsystem.

A turbomachine includes a compressor configured to compress air received at an intake portion to form a compressed airflow that exits into an outlet portion. A combustor is operably connected with the compressor, and receives the compressed airflow. A turbine is operably connected with the combustor, and receives the combustion gas flow. The turbine has a plurality of wheels and a plurality of buckets. The turbine receives compressor bleed off air to cool the wheels and buckets. A cooling system is operatively connected to the turbine. The cooling system includes a plurality of heat pipes located axially upstream of at least one of the wheels. The heat pipes are operatively connected to a bearing cooler system. The heat pipes and the bearing cooler system are configured to transfer heat from the compressor bleed off air to one or more heat exchangers.

A blade for a gas turbine is described, the blade comprising a trailing edge and a trailing edge cooling channel extending from a first upstream end to a second downstream end, in which the width of the trailing edge cooling channel in the direction perpendicular to the camber line of the blade varies and the narrowest width in the trailing edge cooling channel is at the first upstream end, so as to provide a blade where the trailing edge can be removed with a minimal change or no change in the cooling flow capacity through the trailing edge cooling channel. Related methods are also described.

A rotor blade for a gas turbine that includes an airfoil and a tip shroud. The tip shroud may have a seal rail that projects radially from an outboard surface and extends circumferentially. The tip shroud may further include: a rotationally leading circumferential face; a rotationally trailing circumferential face; and an outboard face of the seal rail. The tip shroud may be circumferentially divided into three parallel reference zones: a rotationally leading edge zone, a rotationally trailing edge zone and, formed between and separating those, a middle zone. The seal rail may include a hollow cavity wholly contained within at least one of the rotationally leading edge zone and the rotationally trailing edge zone. The cavity may include a mouth formed through at least one of the rotationally leading circumferential face, the rotationally trailing circumferential face, and the outboard face of the seal rail.