Abstract: A stem (34) extends from a second part (30) through a hole (28) in a first part (22). A groove (38) around the stem provides a non-threaded contact surface (42) for a ring element (44) around the stem. The ring element exerts an inward force against the non-threaded contact surface at an angle that creates axial tension (T) in the stem, pulling the second part against the first part. The ring element is formed of a material that shrinks relative to the stem by sintering. The ring element may include a split collet (44C) that fits partly into the groove, and a compression ring (44E) around the collet. The non-threaded contact surface and a mating distal surface (48) of the ring element may have conic geometries (64). After shrinkage, the ring element is locked onto the stem.

Abstract: A method of joining a first part together with a second part is disclosed. The method may comprise providing a first part having a first joining portion including a first channel and a second part having a second joining portion including a second channel. The method may further comprise positioning the first part adjacent to the second part such that the first channel and second channel align with one another to define a cavity. The method may still further comprise preparing a mixture comprising at least one of a metal powder and a polymer binder, placing the mixture into the cavity so as to form a preform and solidifying the preform forming a metal element in the cavity. The metal element joins the first part together with the second part.

Abstract: A stationary turbine engine component, such as a turbine vane, includes a internal spar and an external skin. The internal spar is made of a plurality of spar laminates, and the external skin is made of a plurality of skin laminates. The plurality of skin laminates interlockingly engage the plurality of spar laminates such that the external skin is located and held in place. This arrangement allows alternative high temperature materials to be used on turbine engine components in areas where their properties are needed without having to make the entire component out of such material. Thus, the manufacturing difficulties associated with making an entire component of such a material and the attendant high costs are avoided. The skin laminates can be made of advanced generation single crystal superalloys, intermetallics and refractory alloys.

Abstract: A modular airfoil assembly (200) and related method for interlocking components of an airfoil structure (210) including a platform (220), an airfoil (210) having a shoulder (230) and a stem (232) extending outward from the shoulder. A ring element (100) positioned against the stem (232) secures the shoulder (230) against the platform (210). First and second members (100a, 100b) of the ring element (100) are bonded together with a portion (128j) of a surface (112a) of the second member (100b) extending within and bonded to a portion (128i) of a surface (112b) of the first member (100a).

Abstract: A turbine airfoil (22A) is formed by a first process using a first material. A platform (30A) is formed by a second process using a second material that may be different from the first material. The platform (30A) is assembled around a shank (23A) of the airfoil. One or more pins (36A) extend from the platform into holes (28) in the shank (23A). The platform may be formed in two portions (32A, 34A) and placed around the shank, enclosing it. The two platform portions may be bonded to each other. Alternately, the platform (30B) may be cast around the shank (23B) using a metal alloy with better castability than that of the blade and shank, which may be specialized for thermal tolerance. The pins (36A-36D) or holes for them do not extend to an outer surface (31) of the platform, avoiding stress concentrations.

Abstract: A method of forming a rotor for a turbine engine such that the rotor is formed of two materials including: an inner disk formed from a first material, such as steel, and an outer ring formed from a second material, such as a nickel alloy, having a larger thermal expansion coefficient than the first material forming the inner disk. The ring may include an inner aperture having a conical shape, and the disk may have an outer surface with a conical shape and a diameter with a portion that is larger than a portion of the ring. The ring may be heated such that the aperture expands to a size greater than the largest diameter of the inner disk. The ring may be positioned over the disk and allowed to cool to allow the ring to be attached to the disk. The ring and disk may then be co-forged.

Abstract: A turbine blade having a squealer tip coupled to a radially outer end of the turbine blade that is usable in a gas turbine engine is disclosed. The squealer tip may require less cooling air and may therefore be more efficient than conventional configurations. The squealer tip may be formed from one or more materials such as oxide dispersion strengthened alloys and FeCrAl alloys. The squealer tip may be formed from a plurality of segmented tips extending radially outward and spaced apart from each other. For example, the squealer tip may be formed from two rails extending radially outward and spaced apart from each other. The two rails may be formed from outer and inner rails that each form a continuous ring. The squealer tip may be attached to the tip with a transient liquid phase bond or an additive manufacturing process, such as, a selective laser melting process.

Abstract: A component for use in a turbine engine including a first member and a second member associated with the first member. The second member includes a plurality of connecting elements extending therefrom. The connecting elements include securing portions at ends thereof that are received in corresponding cavities formed in the first member to attach the second member to the first member. The connecting elements are constructed to space apart a first surface of the second member from a first surface of the first member such that at least one cooling passage is formed between adjacent connecting elements and the first surface of the second member and the first surface of the first member.

Abstract: A method of making a combustion turbine component includes forming a metallic body by direct metal fabrication (DMF) to have at least one surface portion defining a first plurality of surface cooling features each having a first dimension and at least one second surface cooling feature on at least one of the first plurality of surface cooling features and having a second dimension less than said first dimension and less than 200 ?m. Forming the metallic body by DMF may include forming a plurality of metallic combustion turbine subcomponent greenbodies by DMF and assembling the plurality of metallic combustion turbine subcomponent greenbodies together to form a metallic greenbody assembly. The metallic greenbody assembly may be sintered to thereby form the metallic body.

Abstract: Multi-component vane segment and method for forming the same. Assembly includes: positioning a pre-formed airfoil component (12) and a preformed shroud heat resistant material (18) in a mold, wherein the airfoil component (12) and the shroud heat resistant material (18) each comprises an interlocking feature (24); preheating the mold; introducing molten structural material (46) into the mold; and solidifying the molten structural material such that it interlocks the pre-formed airfoil component (12) with respect to the preformed shroud heat resistant material (18) and is effective to provide structural support for the shroud heat resistant material (18). Surfaces between the airfoil component (12) and the structural material (46), between the airfoil component (12) and the shroud heat resistant material (18), and between the shroud heat resistant material (18) and the structural material (46) are free of metallurgical bonds.

Abstract: Bi-casting a platform (50) onto an end portion (42) of a turbine airfoil (31) after forming a coating of a fugitive material (56) on the end portion. After bi-casting the platform, the coating is dissolved and removed to relieve differential thermal shrinkage stress between the airfoil and platform. The thickness of the coating is varied around the end portion in proportion to varying amounts of local differential process shrinkage. The coating may be sprayed (76A, 76B) onto the end portion in opposite directions parallel to a chord line (41) of the airfoil or parallel to a mid-platform length (80) of the platform to form respective layers tapering in thickness from the leading (32) and trailing (34) edges along the suction side (36) of the airfoil.

Abstract: A turbine airfoil usable in a turbine engine and including a depth indicator for determining outer wall blade thickness. The airfoil may include an outer wall having a plurality of grooves in the outer surface of the outer wall. The grooves may have a depth that represents a desired outer surface and wall thickness of the outer wall. The material forming an outer surface of the outer wall may be removed to be flush with an innermost point in each groove, thereby reducing the wall thickness and increasing efficiency. The plurality of grooves may be positioned in a radially outer region of the airfoil proximate to the tip.

Abstract: A combustion turbine component (10) includes a combustion turbine component substrate (16) and an alloy coating (14) on the combustion turbine component substrate. The alloy coating (14) includes nickel (Ni), chromium (Cr), aluminum (Al), and yttrium (Y). Furthermore, the alloy coating includes at least one of titanium (Ti), tantalum (Ta), tungsten (W), and rhenium (Re). The alloy coating also includes at least one rare earth element, and an oxide of at least one of the yttrium (Y) and the at least one rare earth element.

Abstract: A method for making a gas turbine component (100). A central core (20) is positioned to occupy a space that will define a central channel (42), and an outer channel core (30) is positioned spaced apart from the central core (20). A mold (35) is formed around the central core (20) and the outer channel core (30), so that an exterior wall (32) contacts the mold (35). A substrate material, such as a metal alloy (247) in liquid form, is added to the mold (35) to form an internal volume (41) of the component (100). The central core (20) and the outer channel core (30) are removed, and interconnect channels (44) are formed between the thus-formed central channel (42) and the inner portion (49) of the outer channel (62) thus far formed. A preform (55) is placed into the inner portion (49) and may have a desired outer surface (57) shape. An overlay material is applied to form an outer layer (60), thus defining the remainder of the outer channel (62), which is obtained upon removal of the preform (55).

Abstract: A combustion turbine component (10) includes a combustion turbine component substrate (16) and an alloy coating (14) on the combustion turbine component substrate. The alloy coating (14) includes a first amount, by weight percent, of cobalt (Co) and a second amount, by weight percent, of nickel (Ni), the first amount being greater than the second amount. The alloy coating further includes chromium (Cr), aluminum (Al), at least one rare earth element, and an oxide of the at least one rare earth element.

Abstract: A combustion turbine component (10) includes a combustion turbine component substrate (16) and an alloy coating (14) on the combustion turbine component substrate. The alloy coating (14) includes a first amount, by weight percent, of nickel (Ni) and a second amount, by weight percent, of cobalt (Co), the first amount being greater than the second amount. The alloy coating also includes chromium (Cr), aluminum (Al), and yttrium (Y). The alloy coating further includes at least one of titanium (Ti), tantalum (Ta), tungsten (W), and rhenium (Re). Moreover, the alloy coating includes at least one rare earth element, and an oxide of at least one of the yttrium the at least one rare earth element.

Abstract: A turbine airfoil (31) with an end portion (42) that tapers (44) toward the end (43) of the airfoil. A ridge (46) extends around the end portion. It has proximal (66) and distal (67) sides. A shroud platform (50) is bi-cast onto the end portion around the ridge without bonding. Cooling shrinks the platform into compression (62) on the end portion (42) of the airfoil. Gaps between the airfoil and platform are formed using a fugitive material (56) in the bi-casting stage. These gaps are designed in combination with the taper angle (44) to accommodate differential thermal expansion while maintaining a gas seal along the contact surfaces. The taper angle (44) may vary from lesser on the pressure side (36) to greater on the suction side (38) of the airfoil. A collar portion (52) of the platform provides sufficient contact area for connection stability.

Abstract: A method of manufacturing a metallic component includes atomizing, in an inert atmosphere, a metallic liquid having at least one rare-earth element and at least one non rare-earth element to form a metallic powder. A series of heat treating steps are performed on the metallic powder. A first heat treating step is performed in an oxidizing atmosphere, and a second heat treating step is performed in an inert atmosphere. A third heat treating step is performed in a reducing atmosphere to form a metallic power having an increased proportion of rare-earth oxides compared to non rare-earth oxides. The metallic component is formed from the metallic powder having the increased proportion of rare-earth oxides compared to non rare-earth oxides.

Abstract: A system for forming a surface coating on an outer surface of a foam for use with cooling system of turbine engines. The system may include removing filler from the outer surface of the foam to expose a porous structure of the foam, whereby portions of the porous structure extend outwardly from a newly formed outer surface of the filler. A surface layer may be applied to the outer surface of the filler and exposed portions of the porous structure, whereby the surface layer is attached to the porous structure at least in part due to mechanical interaction with the portions of the porous structure extending outwardly from the newly formed outer surface of the filler. The filler material may then be removed from the porous structure.

Abstract: A turbine airfoil (22A) is formed by a first process using a first material. A platform (30A) is formed by a second process using a second material that may be different from the first material. The platform (30A) is assembled around a shank (23A) of the airfoil. One or more pins (36A) extend from the platform into holes (28) in the shank (23A). The platform may be formed in two portions (32A, 34A) and placed around the shank, enclosing it. The two platform portions may be bonded to each other. Alternately, the platform (30B) may be cast around the shank (23B) using a metal alloy with better castability than that of the blade and shank, which may be specialized for thermal tolerance. The pins (36A-36D) or holes for them do not extend to an outer surface (31) of the platform, avoiding stress concentrations.