Abstract: A measurement system for a gas turbine engine is provided. The measurement system comprises a sensor assembly. The measurement system also includes multiple sensors coupled to the sensor assembly. The sensor assembly is configured to be removably inserted within a space defined by a circumferential track embedded within an inner diameter of a casing of the gas turbine engine without having to disassemble the casing.

Abstract: A measurement system for a gas turbine engine is provided. The measurement system comprises a sensor assembly. The measurement system also includes multiple sensors coupled to the sensor assembly. The sensor assembly is configured to be removably inserted within a space defined by a circumferential track embedded within an inner diameter of a casing of the gas turbine engine without having to disassemble the casing.

Abstract: Turbine shrouds for turbine systems are disclosed. The turbine shrouds may include a forward end, an aft end, a first and second side, an outer surface facing a cooling chamber formed between the body and the turbine casing, and an inner surface facing a hot gas flow path for the turbine system. The turbine shroud may also include at least one collection plenum extending within the body between the forward end and the aft end. Additionally, the turbine shroud may include set of cooling passage(s) extending within the body. Each of the cooling passages of the set of cooling passage(s) may include an inlet portion in fluid communication with the cooling chamber, an outlet portion in fluid communication with the at least one collection plenum, and an intermediate portion fluidly coupling the inlet portion and the outlet portion.

Abstract: A blade for a turbomachine includes an airfoil extending radially between a root and a tip with a tip shroud coupled to the tip of the airfoil. The tip shroud includes a platform having an outer surface extending generally perpendicular to the airfoil. The tip shroud also includes a forward rail extending radially outward from the outer surface of the platform. The forward rail is oriented generally perpendicular to a hot gas path of the turbomachine. A cooling cavity is defined in a central portion of the platform. The tip shroud also includes a cooling channel extending between the cooling cavity and an ejection slot formed in the forward rail. The ejection slot is positioned radially outward of the outer surface of the platform of the tip shroud.

Abstract: Turbine shrouds for turbine systems are disclosed. The turbine shrouds may include a forward end, an aft end, a first and second side, an outer surface facing a cooling chamber formed between the body and the turbine casing, and an inner surface facing a hot gas flow path for the turbine system. The turbine shroud may also include at least one collection plenum extending within the body between the forward end and the aft end. Additionally, the turbine shroud may include set of cooling passage(s) extending within the body. Each of the cooling passages of the set of cooling passage(s) may include an inlet portion in fluid communication with the cooling chamber, an outlet portion in fluid communication with the at least one collection plenum, and an intermediate portion fluidly coupling the inlet portion and the outlet portion.

Abstract: A damper pin for damping adjacent turbine blades coupled to a rotor disk comprises a first elongated body having a center portion disposed between a first end portion and a second end portion. The first end portion, center portion and second end portion define a generally arcuate top portion of the first elongated body. The first elongated body defines a slot that extends axially therethrough. The damper pin further includes a second elongated body having a generally arcuate top portion. The second elongated body is at least partially disposed within the slot and is slideably engaged in an axial direction with the first elongated body.

Abstract: A turbine bucket assembly and turbine system are disclosed. The turbine bucket assembly includes a single-lobe joint having an integral platform, the joint having a first axial length; a segmented airfoil having a root segment extending radially outward from the platform and a tip segment coupled to the root segment, the tip segment having a second axial length, which is less than the first axial length; and a turbine wheel defining a receptacle with a geometry corresponding to the single-lobe joint and being coupled to the single-lobe joint. The tip segment includes a tip segment material, the root segment includes a root segment material, and the turbine wheel includes a turbine wheel material, the root segment material and the turbine wheel material having a lower heat resistance and a higher thermal expansion than the tip segment material.

Abstract: Various aspects include a composite turbomachine component and related methods. In some cases, a method includes: identifying a location of potential or actual structural weakness in a body of a turbomachine component, the body including a first material having a first thermal expansion coefficient; forming a slot in the location of the body, the slot extending at least partially through a wall of the turbomachine component; and bonding an insert to the body at the slot to form a composite component, the insert including a second material having a second thermal expansion coefficient differing from the first thermal expansion coefficient by up to approximately ten percent, the second material consisting of a nickel-chromium-molybdenum alloy, wherein after the bonding the insert is configured to reduce the potential or actual structural weakness in the body.

Abstract: A turbine bucket according to various embodiments includes: a base; a blade coupled to the base and extending radially outward from the base, the blade including: a body having: a pressure side; a suction side opposing the pressure side; a leading edge between the pressure side and the suction side; and a trailing edge between the pressure side and the suction side on a side opposing the leading edge; a plurality of radially extending cooling passageways within the body; and at least one bleed aperture fluidly coupled with at least one of the plurality of radially extending cooling passageways, the at least one bleed aperture extending through the body at the trailing edge; and a shroud coupled to the blade radially outboard of the blade.

Abstract: A blade for a turbomachine includes an airfoil extending radially between a root and a tip with a tip shroud coupled to the tip of the airfoil. The tip shroud includes a platform having an outer surface extending generally perpendicular to the airfoil. The tip shroud also includes a forward rail extending radially outward from the outer surface of the platform. The forward rail is oriented generally perpendicular to a hot gas path of the turbomachine. A cooling cavity is defined in a central portion of the platform. The tip shroud also includes a cooling channel extending between the cooling cavity and an ejection slot formed in the forward rail. The ejection slot is positioned radially outward of the outer surface of the platform of the tip shroud.

Abstract: A turbine nozzle includes an airfoil that extends in span from an inner band to an outer band where the inner band and the outer band define inner and outer flow boundaries of the turbine nozzle. At least one of the inner band and the outer band defines a plurality of cooling channels formed and a coolant discharge plenum beneath a gas side surface of the corresponding inner or outer band that is in fluid communication with the cooling channels. The coolant discharge plenum is formed within the inner band or the outer band downstream from the cooling channels and upstream from a plurality of coolant discharge ports.

Abstract: Various embodiments include a system having: a computing device configured to detect deformation in a manufactured component by: obtaining a post-deployment three-dimensional (3D) depiction of the manufactured component; obtaining a model of the manufactured component including: a nominal shape model indicating a nominal shape of the manufactured component prior to operational deployment, and an expected deformation model indicating expected deformation of the manufactured component after operational deployment; aligning a localized region of the manufactured component in the post-deployment 3D depiction with the localized region of the manufactured component in the nominal shape model; identifying a first set of points in the localized region not subject to deformation between the post-deployment 3D depiction and the nominal shape model; and identifying a second set of points in the localized region subject to deformation between the post-deployment 3D depiction and the nominal shape model.

Abstract: This disclosure provides systems and methods for sealing flow path components, such as turbomachine airfoils, with a front-leaded seal. A seal channel is defined between a portion of the suction side surface of a first flow path component and a portion of the pressure side of a second flow path component. A seal is retained within the seal channel formed by the pressure side portion and the suction side portion and the seal channel defines a forward opening through which the seal is inserted during installation.

Abstract: This disclosure provides systems and methods for sealing flow path components, such as turbomachine airfoils, with a front-leaded seal. A seal channel is defined between a portion of the suction side surface of a first flow path component and a portion of the pressure side of a second flow path component. A seal is retained within the seal channel formed by the pressure side portion and the suction side portion and the seal channel defines a forward opening through which the seal is inserted during installation.

Abstract: A turbine bucket includes a leading edge, a trailing edge, a root portion, and a tip portion. The turbine bucket also includes one or more cooling passages extending through a body of the turbine bucket from an inlet to an outlet. The cooling passages are configured to route a cooling flow of fluid through the turbine bucket. The turbine bucket further includes a plenum defined within the tip portion to receive the fluid from the outlet of the cooling passages for expulsion of the cooling flow of fluid into a main flow path via at least one outlet hole proximate the trailing edge of the turbine bucket.

Abstract: A turbine bucket according to various embodiments includes: a base; a blade coupled to the base and extending radially outward from the base, the blade including: a body having: a pressure side; a suction side opposing the pressure side; a leading edge between the pressure side and the suction side; and a trailing edge between the pressure side and the suction side on a side opposing the leading edge; a plurality of radially extending cooling passageways within the body; and at least one bleed aperture fluidly coupled with at least one of the plurality of radially extending cooling passageways, the at least one bleed aperture extending through the body at the trailing edge; and a shroud coupled to the blade radially outboard of the blade.

Abstract: A turbine nozzle includes an airfoil that extends in span from an inner band to an outer band where the inner band and the outer band define inner and outer flow boundaries of the turbine nozzle. At least one of the inner band and the outer band defines a plurality of cooling channels formed and a coolant discharge plenum beneath a gas side surface of the corresponding inner or outer band that is in fluid communication with the cooling channels. The coolant discharge plenum is formed within the inner band or the outer band downstream from the cooling channels and upstream from a plurality of coolant discharge ports.

Abstract: A damper pin for damping adjacent turbine blades coupled to a rotor disk comprises a first elongated body having a center portion disposed between a first end portion and a second end portion. The first end portion, center portion and second end portion define a generally arcuate top portion of the first elongated body. The first elongated body defines a slot that extends axially therethrough. The damper pin further includes a second elongated body having a generally arcuate top portion. The second elongated body is at least partially disposed within the slot and is slideably engaged in an axial direction with the first elongated body.

Abstract: A turbine blade having an airfoil defined by a concave shaped pressure side outer wall and a convex shaped suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber for receiving the flow of a coolant. The turbine blade may include a rib configuration that partitions the chamber and defines a flow passage having a first and a second side. The flow passage may include a port formed through the first side. A projection of a center axis of the port through the flow passage may define a strike point on the second side of the flow passage, and a backstrike recess may be positioned on the second side of the flow passage so to contain the strike point.

Abstract: A turbomachine includes a compressor portion, and a turbine portion operatively connected to the compressor portion. The turbine portion includes a compressor discharge portion, a hot gas path and a first stage nozzle arranged at the hot gas path. A combustor assembly is fluidically connected to the compressor portion. A transition piece has an inlet portion fluidically connected to the combustor assembly and an outlet portion fluidically connected to the turbine portion. A seal member is operatively connected relative to one of the transition piece and the first stage nozzle of the turbine portion. The seal member extends from a first end to a second, cantilevered end through an intermediate portion including one or more openings that provide a purge fluid flow path from the compressor discharge portion toward the hot gas path.