Abstract: A method is provided for drilling a hole through a metallic workpiece having a thermal barrier coating with a ceramic top coat by laser drilling a counterbore to a depth which extends through the ceramic top coat but not substantially into the metallic workpiece and then laser drilling the hole through the workpiece aligned with the counterbore, the counterbore having a diameter larger than the hole.

Abstract: A turbine blade for a gas turbine engine. An existing blade was found to exhibit bowing, or a concave configuration facing the pressure side, along its trailing edge. The invention reduces bowing by (1) changing tilt, (2) changing lean, (3) reducing the number of cooling holes, while (4) changing the diameters of the cooling holes, to maintaining the total cooling flow unchanged.

Abstract: A turbine blade for a gas turbine engine. An existing blade was found to exhibit bowing, or a concave configuration facing the pressure side, along its trailing edge. The invention reduces bowing by (1) changing tilt, (2) changing lean, (3) reducing the number of cooling holes, while (4) changing the diameters of the cooling holes, to maintaining the total cooling flow unchanged.

Abstract: An airfoil for a gas turbine engine, the airfoil having a core body with an airfoil body, an integral partial height squealer tip defining a tip shelf, and an integral tip cap between the airfoil body and the integral partial height squealer tip; and a squealer tip extension bonded to the partial height squealer tip. A method for manufacture and repair of an airfoil.

Abstract: A gas turbine engine includes rotor blades including airfoils that facilitates reducing manufacturing losses due to airfoil trailing edge scarfing. Each airfoil includes a first and second sidewall connected at a leading edge and a trailing edge. The sidewalls define a cooling cavity that includes at least a leading edge chamber bounded by the sidewalls and the airfoil leading edge, and a trailing edge chamber bounded by sidewalls and the airfoil trailing edge. The cooling cavity trailing edge chamber includes a tip region, a throat, and a passageway region connected in flow communication such that the throat is between the tip region and the passageway region. Furthermore, the tip region is bounded by the airfoil tip and extends divergently from the throat, such that a width of the tip region is greater than a width of the throat.

Abstract: A method of manufacturing a gas turbine engine fan blade (10) comprises forming three metal workpieces (30,32,34). The metal workpieces (30,32,34) are assembled into a stack (36) so that the flat surfaces (38,42,46,48) are in mating abutment. Heat and pressure is applied across the thickness of the metal workpieces (30,32,34) to diffusion bond the metal workpieces (30,32,34) together to form an integral structure (80). The integral structure (80) is upset forged at one end (58) to produce an increase in thickness (82) for forming the blade root (26). The upset forged integral structure (80) is then hot creep formed and superplastically formed to produce the required aerofoil shape and the thickened end (82) is machined to form the blade root (26). The method enables thinner metallic workpieces with better microstructure to be used and increases the yield of metallic workpieces.

Abstract: The present invention relates to a turbine blade or vane (13; 14), in particular for a gas turbine (10), which has at least one chamber (18; 29; 20; 21) which can be acted on by a cooling medium and, at a rear edge (26), has a gap (25), which is delimited by two walls (28, 29), for the cooling medium to be discharged. According to the invention, at least one of the walls (28; 29) can be remachined in order to change the cross section (A) of the gap (25). In this way, the cross section (A) can be adapted to the particular boundary conditions, and therefore the consumption of cooling medium can be minimized. The invention also relates to a process for producing a turbine blade or vane (13; 14) of this type.

Abstract: A core for use in casting a turbine bucket including serpentine cooling passages is divided into two pieces including a leading edge core section and a trailing edge core section. Wall thicknesses at the leading edge and the trailing edge of the turbine bucket can be controlled independent of each other by separately positioning the leading edge core section and the trailing edge core section in the casting die. The controlled leading and trailing edge thicknesses can thus be optimized for efficient cooling, resulting in more efficient turbine operation.

Abstract: The cooling scheme of a gas turbine engine component is modified to improve local cooling without redesigning the investment casting. The modification includes forming at least one channel in the component such that the channel is in fluid communication with a cooling medium source associated with the component. The channel is then partially filled with a removable material, and the removable material is covered with a patch material so as to completely fill the channel. Lastly, the removable material is removed from the channel so as to create an internal cooling passage in the component that is in fluid communication with the cooling medium source.

Abstract: Method for producing or repairing cooling channels (7a) of a gas turbine component (1), whereby the cooling channels (7a) are masked with a thermally stable filling material (3) on the cast gas turbine component (1), and another epitactic layer (6) is created above the ceramic material (3) with the help of a laser (4) and a powder (5). The thermally stable filling material (3) is removed by etching.

Abstract: A cooled component, such as a turbine blade for gas turbines is provided, having efficient internal cooling, with an interior cooling passageway having a round cross-section. A row of feeding holes for the coolant are arranged spaced from each other in the direction of the longitudinal axis of the cooling passageway and originating from a common coolant channel. Each of the feeding holes intersects the cooling passageway tangentially. The ease of manufacturing the cooling component is improved in that the majority of the feeding holes have a hole diameter that is smaller than half of the hydraulic diameter of the cooling passageway, and selected feeding holes have a hole diameter that is greater than half of the hydraulic diameter of the cooling channel.

Abstract: A method for transpiration cooling of the flow path surface region of an engine component used in a gas turbine engine by channeling a substrate to provide cooling channels through the substrate having a diameter of about 0.005″ to about 0.02″ to allow passage of cooling fluid from a cooling fluid source through the cooling channels. A bond coat of about 0.0005″ to about 0.005″ in thickness is applied to the substrate such that the bond coat partially fills the cooling channels. A porous TBC of at least about 0.003″ to about 0.01″ thick is applied over the bond coat, such that the TBC completely fills the cooling channels. Cooling fluid from a cooling fluid source is passed through the cooling channels and porous TBC. Because the channel exit is filled with TBC, the cooling fluid is transmitted through the porous passageways of the TBC. The porous passageways provide a plurality of tortuous routes to the TBC surface.

Abstract: A turbine nozzle includes a plurality of airfoils and integral hubs brazed into corresponding seats of outer and inner bands. The hubs have brazeless fillets blending the bands to the airfoils.

Abstract: A process for manufacturing components of gas turbine engine combustors, including the steps of performing a plurality of forming steps on a workpiece of a designated material to produce a combustor component of desired size and shape without exceeding a forming limit of the designated material, the forming steps being performed without any intermediate heat treatment on the workpiece.

Abstract: A method for accurately sizing and forming the cooling holes of an air-cooled gas turbine engine component on which a protective diffusion coating will be deposited. The method generally entails drilling a hole in a surface region of a substrate, and then measuring the thickness of any recast surface region surrounding the hole and created as a result of a portion of the surface region having melted and then resolidified. The thickness of the additive layer of the diffusion coating that will deposit on a corresponding recast surface region of the component is then predicted based on an inverse relationship determined to exist with the thickness of the recast surface region formed during drilling of the cooling hole. An appropriately-oversized hole can then be formed in the component so that, after depositing the diffusion coating on the component, the additive layer grows sufficiently within the hole to yield a cooling hole approximately having the required final diameter.

Abstract: A hybrid gas turbine engine airfoil is made by forming a metal airfoil with a pocket in one side thereof. The pocket is covered by a caul. An elastomeric fluid is injected into the pocket and cured therein for bonding thereto. The caul is removed from the airfoil to expose the cured elastomer conforming to the aerodynamic profile thereof.

Abstract: Process of plugging cooling holes (4) of a gas turbine component (1) with an external surface (7), having a cavity (2) and a plurality of cooling holes (4) before coating the gas turbine component (4), characterised in that mask material (6) is applied to the external surface (7) of the component (1) while a fluid is forced through the cooling holes (4), the mask material (6) is thickened while the fluid is forced through the cooling holes (4), the cooling holes (4) are plugged with plug material (6b), the mask material (6a) is removed from the external surface (7) of the component (1), the component (1) and the plugged cooling holes (4) are coated and the plug material (6b) is removed from the cooling holes (4).

Abstract: A method for cooling the trailing edge 126, 146 region of a stator vane platform 48, 54 and for forming a cooling passage 198, 214 using platform slots in the side of a stator vane is disclosed. Various steps are developed which provide for effective cooling with minimal intrusion into the working medium flowpath 18. In one particular embodiment, the method includes extending a feather seal slot rearwardly with an extension 88, 96 in facing sides 116, 118, or 136, 138 of a pair of vane platforms 48, 54 and for flowing cooling fluid laterally prior to discharging the cooling air for impingement cooling of an adjacent stator vane.

Abstract: A method for modifying cooling holes in a gas turbine engine film-cooled component by machining cooling hole outlets to enlarge the outlets and remove any portion of the cooling hole walls which might exhibit cracks.

Abstract: A turbine engine having improved high pressure turbine cooling is disclosed. In the engine, relatively cool intermediate pressure (P2x) air is diverted from a region of a compressor section and over a high work turbine blade at a lower static pressure than the diverted air to cool the blade. Advantageously, as the diverted air is relatively cool, use of a conventional TOBI nozzle may be eliminated. Similarly, showerheads on the blade may be eliminated. As well, the diverted air may conveniently be used to seal a rear bearing compartment within the engine.

Abstract: A method of forming an internal channel within an article, such as a cooling channel in an air-cooled blade, vane, shroud, combustor or duct of a gas turbine engine. The method generally entails forming a substrate to have a groove recessed in its surface. A solid member is then placed in the groove, with the solid member being sized and configured to only partially fill the groove so that a void remains in the groove. The void is then filled with a particulate material so that the groove is completely filled. A layer is then deposited on the surface of the substrate and over the solid member and the particulate material in the groove, after which at least the solid member is removed from the groove to form the channel in the substrate beneath the layer.

Abstract: A hollow airfoil is provided which includes a body, a trench, and a plurality of cooling apertures disposed within the trench. The body extends chordwise between a leading edge and a trailing edge, and spanwise between an outer radial surface and an inner radial surface, and includes an external wall surrounding a cavity. The trench is disposed in the external wall along the leading edge, extends in a spanwise direction, and is aligned with a stagnation line extending along the leading edge.