Abstract: A turbine engine blade outer air seal segment has a body having a base portion. The base portion has a transversely concave ID face, a forward end, an aft end, and first and second circumferential edges. The body has at least one mounting hook. At least one cover plate is secured to the body to define at least one cavity. The cover plate has a plurality of feed holes. A plurality of outlet holes extend through the base portion to the ID face. At least one of the base portion and cover plate comprises a protruding portion protruding into the cavity to form a partial restriction separating circumferentially and fore-aft offset cavity portions.

Abstract: A turbine blade is cooled by cooling air that flows through a radial cooling channel. The turbine blade includes a root and an airfoil. The flow of cooling air into the cooling channel is limited by a pre-meter orifice to provide a reduced pressure within the cooling channel. The pressure drop results from the cross-sectional area of the pre-meter orifice being less than the cross-sectional area of the adjacent cooling channel. After flowing through the cooling channel, the cooling air exits the cooling channel through a film hole to form a film layer over the airfoil to cool and insulate the turbine blade.

Abstract: A cooling system for an airfoil portion of a turbine engine component is provided. The cooling system includes a first cavity dedicated to cooling a trailing edge portion of an airfoil portion and a second cavity dedicated to cooling an aft portion of a pressure side wall.

Abstract: A turbine component includes a coolant passage and a plurality of trip members that extend within the coolant passage. Each of the plurality of trip members includes an orientation relative to an expected flow direction through the coolant passage. At least three of the trip members have orientations that are different from each other. For example, the orientations of the trip members progressively transition along the length of the coolant passage.

Abstract: A turbine engine component has an airfoil portion with a pressure side wall and a suction side wall and a cooling system. The cooling system has at least one cooling circuit disposed longitudinally along the airfoil portion. Each cooling circuit has a plurality of staggered internal pedestals for increasing heat pick-up.

Abstract: A method is disclosed that selectively applies thermal barrier coatings that exhibit different degrees of thermal conductivity to different inner surface areas of engine combustor panels. Different types of TBCs are applied to predetermined inner surface areas of a combustor panel based on empirical observation or prediction. TBCs exhibiting low thermal conductivity are applied to combustor panel areas that are exposed to hotter temperatures and TBCs exhibiting higher thermal conductivity are applied to areas that are exposed to lower temperatures.

Abstract: A method for manufacturing a turbine engine component comprises the steps of forming a first half of an airfoil portion of the turbine engine component and forming a plurality of microcircuits having at least one passageway on an exposed internal wall of the first half of the airfoil portion. The method further comprises forming a second half of the airfoil portion of said turbine engine component, and forming at least one additional cooling microcircuit having at least one passageway on an exposed internal wall of the second half of the airfoil portion. Thereafter, the first half is placed in an abutting relationship with the second half after the cooling microcircuits have been formed and inspected. The first half and the second half are joined together to form the airfoil portion.

Abstract: A turbine blade cooling system for a gas turbine engine includes a turbine blade having a trailing edge, a concave side, and a convex side. The trailing edge defines at least one set of impingement holes each having a central longitudinal axis which is closer to a nearest portion of an edge of the blade at one of the concave and convex sides relative to a nearest portion of an edge of the blade at the other of the concave and convex sides. Alternatively or in addition to the foregoing, the trailing edge can define at least one set of impingement holes each having a central longitudinal axis which is angled in a direction of a flow of cooling medium toward one of the concave and convex sides relative to the other of the concave and convex sides.

Abstract: A gas turbine engine is provided with a dirt separator placed in a path of the radially outer cooling air flow, and includes a radially outer leg that defines a space to capture dirt or other impurities. An access port is provided in a turbine engine case in an area adjacent to the dirt separator. The access port has two openings at a tangent to the centerline of the gas turbine engine covered by access port covers which can be removed to allow a cleaning tool to have access through the access port.

Abstract: A method for forming a component for use in a gas turbine engine, such as a turbine vane construction is provided. The method broadly comprises the steps of: forming a first aerodynamic structure having a first platform with a leading edge and a trailing edge, and an edge with an airfoil suction side structure; forming a second aerodynamic structure having a second platform with a leading edge and a trailing edge, and an first edge with an airfoil pressure side structure; and joining the two structures together so that the airfoil suction side structure mates with the airfoil pressure side structure to form an airfoil.

Abstract: A turbine engine structure is provided that includes a wall having an exterior surface defining an internal passage. A locating hole extends through the wall from the exterior surface to the passage. A film hole is recessed in the exterior surface and adjoins the locating hole. The film hole and locating hole are in communication with the passage. The locating hole is formed during the casting process in which a core is supported with a locating pin. Upon removal of the locating pin, the locating hole is formed. The locating holes can be used to determine a position of features of the structure for subsequent processing operations of the structure. For example, film holes are machined in the exterior surface, such as by an electrical discharge machining process, to intersect the locating holes.

Abstract: A gas turbine engine is provided with a turbine rotor disk having a plurality of disk slots which each receive a turbine blade. Cooling air is delivered into the disk slots, and flows into cooling air passages at a radially inner surface of the turbine blades. As known, the cooling air flows through the turbine blades to cool internal and external surfaces of the turbine blade. The disk slots are formed to have a greater cross-sectional area adjacent at the inlet end than at downstream locations. In this manner, the velocity of the air at the inlet end is reduced compared to the prior art, and pressure losses are minimized compared to the prior art.

Abstract: A method for manufacturing a turbine engine component comprises the steps of forming a first half of an airfoil portion of the turbine engine component and forming a first cooling microcircuit having at least one passageway on an exposed internal wall of the first half of the airfoil portion. The method further comprises forming a second half of the airfoil portion of said turbine engine component, forming a second cooling microcircuit having at least one passageway on an exposed internal wall of the second half of the airfoil portion, and placing the first half in an abutting relationship with the second half after the cooling microcircuits have been formed and inspected.

Abstract: A gas turbine engine is provided with a radially outer cooling air flow. A dirt separator is placed in the path of the radially outer cooling air flow, and includes a radially outer leg that defines a space to capture dirt or other impurities. A radially inner leg of the dirt separator includes open air flow passages to allow air to flow through the inner leg and downstream to cool various components within the gas turbine engine.

Abstract: A gas turbine engine blade has a relatively large fillet to improve the characteristics of the air flow thereover. The fillet has a thin wall which, together with an impingement rib, defines a fillet cavity therebetween, and cooling air is provided to flow through impingement holes in the impingement rib and impinge on the rear surface of the fillet. The impingement holes are elongated in cross sectional shape with their elongations being orient in a direction generally transverse to a radial direction.

Abstract: A turbine blade airfoil assembly includes a cooling air passage. The cooling air passage includes a plurality of impingement openings that are isolated from at least one adjacent impingement opening. The cooling air passage is formed and cast within a turbine blade assembly through the use of a single core. The single core forms the features required to fabricate the various separate and isolated impingement openings. The isolation and combination of impingement openings provides for the augmentation of convection and film cooling and provide the flexibility to tailor airflow on an airfoil to optimize thermal performance of an airfoil.

Abstract: An internally cooled gas turbine engine turbine vane has an outboard shroud and an airfoil extending from an outboard end at the shroud to an inboard end. A cooling passageway has an inlet in the shroud, a first turn at least partially within the airfoil, a first leg extending from the inlet inboard through the airfoil to the first turn, and a second leg extending from the first turn. A dividing wall is in the passageway and has an upstream end in an outboard half of a span of the airfoil and has a plurality of vents. The vane may be formed as a reengineering of a baseline configuration lacking the dividing wall.

Abstract: An airfoil, and in a disclosed embodiment a rotor blade, has a serpentine cooling path. To best account for the Coriolis effect, the paths of the serpentine cooling channel have trapezoidal cross-sections. An area of the rotor blade between a smaller side of the trapezoidal-shaped paths, and a facing wall of the rotor blade has high thermal and mechanical stresses, and is a challenge to adequately cool. A microcircuit, which is a very thin cooling circuit having crossing pedestals, is embedded into the blade in this area. The microcircuit provides additional cooling, and addresses the challenges with regard to cooling these areas.

Abstract: A shroud includes a panel having leading and trailing edges and inner and outer surfaces extending therebetween. A cooling passage extends through the panel from adjacent the leading edge to an intermediate location and is generally parallel to the inner and outer surfaces. The cooling passage has an inlet for receiving cooling air from adjacent the panel outer surface, a plurality of spaced apart turbulators disposed adjacent to the panel inner surface, and an outlet disposed at the leading edge for discharging the cooling air from the passage.

Abstract: A gas turbine engine rotor blade includes an airfoil having first and second sidewalls joined together at leading and trailing edges. First and second tip walls extend from adjacent the leading edge along the respective first and second sidewalls to adjacent the trailing edge and are spaced apart to define a tip cavity therebetween. A first notch is disposed in the first tip wall adjacent to the leading edge for channeling into the tip cavity a portion of combustion gases flowable over the airfoil for reducing the heating effect of the gases on the blade tip. In a preferred embodiment, a second notch is disposed adjacent to the trailing edge for promoting flow through the tip cavity from the first notch.