Disclosed are a blade ring segment for a turbine section, a turbine section having the blade ring segment, and a gas turbine having the turbine section. Multiple blade ring segments is installed in a turbine casing accommodating turbine blades rotated by combustion gas from a combustor. The blade ring segment includes an inner panel provided in the turbine casing and having multiple air holes through which cooling air fed from the outside of the turbine casing flows, an outer panel disposed on one side of the inner panel, and a cooling structure protruding from one side of the outer panel so as to form a flowing channel in a zigzag pattern so that cooling air fed through the air holes flows therethrough.

A gas turbine engine includes a compressor section having a first portion and an aft portion. A compressor case clearance (CCC) control system is configured to adjust an amount of bleed air delivered to the front portion and the aft portion based on an in-flight phase of an aircraft. In response to invoking a first mode, the CCC control system delivers air to both the front portion and the aft portion. In response to invoking a second mode, the CCC control system reduces the amount of air delivered to the aft portion prior to transitioning from the cruise phase to the descent phase. Accordingly, clearance areas within the compressor section can be selectively increased during specific portions of the flight to avoid contact between blade tips and the engine case.

A nacelle for a gas turbine engine includes at least one hinge axis extending along the nacelle. At least one cowl is mounted to the nacelle, along the at least one hinge axis, by a plurality of hinges. The plurality of hinges includes at least one latch. The at least one latch includes a clevis and a tang. The clevis and the tang of the at least one latch are configured to receive a hinge pin, through a pin aperture, when the at least one cowl is in a closed condition.

A component for a turbine engine includes a body having an outer surface confronting a combustion air flow path and defining an interior, as well as a first cooling passage having at least a portion supplying cooling air to the interior of the body. The component also includes a cyclone separator having a cooling air inlet, a clean air outlet, and a dirty air outlet.

A turbocharger has a meridionally divided turbine housing defining a first scroll and a first nozzle, and a second scroll and a second nozzle. The first and second nozzles are divided from each other by a shroud plate mounted within the nozzle, for isolating the exhaust gas streams flowing through the two nozzles from each other. A plurality of circumferentially spaced vanes are rotatably mounted at a radially inner side of the nozzles. Leading-edge portions of the vanes are slotted to receive the shroud plate when the vanes are pivoted open. Trailing-edge portions of the vanes are unslotted. A radially inner periphery of the shroud plate includes cutouts for abutting the trailing-edge portions of the vanes in the closed position, and includes contour portions for influencing a flow separation ratio of the turbine.

An apparatus and method an airfoil for a turbine engine, the airfoil comprising an outer wall defining an interior bound by a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The airfoil further includes a first cooling circuit having a first portion and a first tip portion and a second cooling circuit having a second portion and a second tip portion.

A method is provided for manufacturing a composite body. A first longitudinal portion of a continuous length of prepreg material is shaped using a first die of a compression molding tool to provide a shaped first longitudinal portion of the continuous length of prepreg material. The shaped first longitudinal portion is consolidated using the compression molding tool to provide a first portion of the composite body. The first portion of the composite body is moved out of the compression molding tool. The first die is swapped for a replacement first die. A second longitudinal portion of the continuous length of prepreg material is shaped using the replacement first die of the compression molding tool to provide a shaped second longitudinal portion of the continuous length of prepreg material. The shaped second longitudinal portion is consolidated using the compression molding tool to provide a second portion of the composite body.

A stator of a gas turbine engine is provided. A static shroud of the stator defines a platform having a radially outer surface. An impingement plate has impingement holes extending therethrough, and the impingement plate is radially spaced apart from the radially outer surface of the platform. The impingement holes are configured to direct a high-speed cooling air flow transversally to the radially outer surface of the stator shroud. A plurality of protrusions project away from the radially outer surface of the platform along a protrusion axis. The protrusions have a cruciate cross-sectional shape when viewed in a plane normal to the protrusion axis. A method of cooling a stator of a gas turbine engine is also provided.

A rotatable component for a turbomachine and a method for regulating a circumferential ingress of a fluid at a trailing edge of the rotatable component are disclosed. The rotatable component includes an airfoil and a mechanical component. The airfoil includes a pressure side and a suction side. The mechanical component is coupled to the airfoil and includes a forward overhanging portion, and an aft overhanging portion. The forward overhanging portion is disposed at a leading edge of the airfoil and extends longitudinally beyond the leading edge. The aft overhanging portion is disposed at a trailing edge of the airfoil and extends longitudinally beyond the trailing edge, where both the forward and aft overhanging portions further extend circumferentially along the pressure side and the suction side of the airfoil. The aft overhanging portion includes a non-axisymmetric profile for regulating the circumferential ingress of the fluid from the pressure to suction sides.

A stator of a gas turbine engine is provided. A static shroud of the stator defines a platform having a radially outer surface. An impingement plate has impingement holes extending therethrough, and the impingement plate is radially spaced apart from the radially outer surface of the platform. The impingement holes are configured to direct a high-speed cooling air flow transversally to the radially outer surface of the stator shroud. A plurality of protrusions project away from the radially outer surface of the platform along a protrusion axis. The protrusions have a cruciate cross-sectional shape when viewed in a plane normal to the protrusion axis. A method of cooling a stator of a gas turbine engine is also provided.