Abstract: A method of using a masking device to mask a portion of a gas turbine engine component for an electroplating process is provided. The masking device includes a main body having sidewalls, a removeable coverplate having an end plate and one or more locking tabs, a fastener located at least partially within the main body, and a shank engaged with the fastener. The method includes placing the gas turbine engine component through an opening in the main body such that the component is held within the main body by a retention slot. The method includes sliding the one or more locking tabs of the removeable coverplate through the opening in the main body such that the removeable coverplate covers the opening of the main body. The method includes securing the one or more locking tabs in one or more relief slots in the sidewalls of the main body.

Abstract: Systems and methods for repairing compressor blades using hybrid manufacturing techniques. A tool for repairing a compressor blade has a plurality of support bases, and a slot configured for securement of said compressor blade to said mounting tool. The tool has a pair of support bases, each of which is configured for the building of a sacrificial support thereon to stabilize the compressor blade during repair. The tool is coupled to a hybrid manufacturing machinery and the sacrificial supports stabilize the compressor blade during additive and subtractive manufacturing processes.

Abstract: An apparatus and method for producing a core for use in casting a gas turbine component is provided. The apparatus comprises a system having a cavity block with an upper portion, a lower portion, and a recessed cavity in each of the upper portion and the lower portion. Positioned within the cavity block is an adapter insert, and within the adapter insert is positioned a core die insert, where the core die insert is fabricated from a sacrificial material and has a hollow internal profile corresponding to an external surface of a core. A ceramic-based material is supplied to the core die insert where it solidifies. The core die insert is placed into a water-based solution and the core die insert is removed from around the core.

Abstract: The present disclosure provides for improvements in fixturing parts in preparation for an electroplating process, thus reducing part handling. The system provides a reusable masking tool comprising a main body having an opening with a removeable coverplate positioned within the opening and one or more locking tabs engaging a corresponding relief slot in at least one of the sidewalls of the main body. The system also comprises a fastener moveably secured within the main body and extending through a top surface of the main body. The main body, fastener, and coverplate may be fabricated from a polymer material by way of an additive manufacturing process. A shank, fabricated from a conductive material, extends through the fastener and is engaged with the fastener such that upon rotation of the fastener, the shank is drawn into contact with the gas turbine engine component, thus providing a conduit for the electroplating process.

Abstract: A system for directing cooling air into a gas turbine combustor is provided. The system comprises a transition duct coupled to a flow sleeve, where air to be used for combustor cooling and in the combustion process enters a bellmouth of the transition duct, passes through a plurality of struts within the bellmouth, and is distributed to a passage located between the combustion liner and flow sleeve.

Abstract: A system and method for cooling a turbine blade tip shroud is provided. The turbine blade comprises a blade attachment, a platform extending radially outward from the attachment, an airfoil extending radially outward from the platform, and a tip shroud extending radially outward from the airfoil. The tip shroud has one or more knife edges extending radially outward from an outer surface of the tip shroud. One or more cooling passages extend through the airfoil and to the tip shroud. The turbine blade also includes one or more tip plates secured at or near the outer surface of the tip shroud thereby forming a plenum between the outer surface and the one or more tip plates. The one or more tip plates also include a plurality of cooling holes for flowing cooling air through the plenum to cool the tip shroud.

Abstract: A nickel-based alloy is disclosed which is suitable for casting gas turbine components having improved strength and comparative lower density while utilizing commercially available heat treatment cycles. The nickel-based alloy is suitable for providing equiaxed, directionally solidified, and single crystal castings. Methods of providing a cast article of the nickel-based alloy and subjecting the article to heat treatment cycles are also disclosed.

Abstract: A heat transfer mechanism is provided comprising a plurality of turbulators located along a surface of a body, such as a combustion liner. The turbulators have a first side with a first ramp angle, a second side with a second ramp angle, a height, and a base width, where the base width is a function of the height and where the turbulators are spaced an axial distance apart that is a function of the turbulator height.

Abstract: The present disclosure provides a system and method of repairing a flange of an engine case having a damaged section of attachment tabs and/or flange region. One or more damaged attachments tabs extending away from a flange are removed along with a first portion of the flange, thus leaving a second portion of the flange. A replacement component is produced comprising a replacement flange portion and one or more replacement tabs. The replacement flange portion is placed in contact with the second portion of the flange and the replacement flange portion is secured to the second portion of the flange.

Abstract: An apparatus and method for mounting a combustion liner within a flow sleeve of a gas turbine combustion system is disclosed. A mounting system comprises a plurality of low-profile mounting tabs secured to a combustion liner where each of the mounting tabs are placed within slots of flow sleeve pegs when the combustion liner is installed in a flow sleeve. A plurality of liner stop brackets are removably secured to a flange of the flow sleeve and have an arm extending to be adjacent to a top contact surface of the mounting tabs. The mounting system reduces blockage to the surrounding airflow.

Abstract: A tool for reforming a core pattern. A method of reforming a core pattern comprises opening a core reformer tool. The core reformer tool has a first portion and a second portion facing the first portion. The first portion includes a concave member and a first adjustable pin. The second portion includes a convex member and a second adjustable pin. The method includes positioning the core pattern in the core reformer tool and closing the core reformer tool. The method comprises adjusting at least one of the first adjustable pin and the second adjustable pin to alter a surface of the core pattern. The method includes directing cooling air through the core reformer tool to solidify the core pattern and opening the core reformer tool to remove the core pattern.

Abstract: An apparatus and method for reducing the vibrations to a transition duct and improving the assembly and fit-up of a transition duct relative to a combustion system are disclosed. A damping member is provided between the transition duct and its mounting points to the gas turbine engine. The damping member comprises one or more layers of a composite material positioned between layers of sheet metal. An adjustable mounting system is also provided for the transition duct to allow for movement of a transition duct relative to the combustion liner.

Abstract: A system and method for improving sealing at a turbine blade tip shroud, while reducing weight associated with the improved sealing is disclosed. The gas turbine blade incorporates a tip shroud having one or more pockets located therein, where the one or more pockets remove weight from the shroud, thus reducing load on the blade attachment generated by additional sealing at the turbine blade shroud. Methods for incorporating the one or more tip shrouds in a new turbine blade or a repaired turbine blade are also disclosed.

Abstract: Systems and methods for modal testing of blades. A method for modal testing of a blade comprises providing a generally spherical fixture for retaining the blade in a fixed-free configuration. The method includes using an excitation device for exciting the blade while the blade is retained within a blade retaining portion of the fixture. The method comprises using a measurement device to measure a response of the blade to the excitation. The fixture includes an attachment portion configured for the securement of the excitation device to the fixture.

Abstract: A guide vane for use in a compressor discharge plenum of a gas turbine engine is disclosed. The guide vane comprises a guide support system, a first side panel, a second side panel, and a deflector panel secured to the mounting system and extending between the first side panel and the second side panel. The deflector panel has a curved profile substantially in accordance with Cartesian coordinate values of X, Y, and Z as set forth in Table 1 where the X, Y, and Z values are in inches from a center point of a bottom surface of the deflector panel. The coordinate values are connected by smooth continuing arcs to form profile sections and the profile sections are joined together smoothly to form the curved profile of the guide vane.

Abstract: A system for retaining a cross fire tube in a multi-combustor gas turbine engine is disclosed. The system comprises a flow sleeve having a generally annular body, and a flange at a forward end thereof and having one or more recessed portions. A cross fire tube extends through one or more openings in the flow sleeve and is secured in place by a retention clip. The retention clip includes a plurality of fingers which engage the cross fire tube and a mounting plate engaging the one or more recessed portions of the flow sleeve flange so as to create a clip engagement having a lower profile than prior art configurations.

Abstract: A core pattern reformer system for adjusting and setting a core pattern for use in casting a gas turbine blade has a first portion having a first internal face with a concave portion. The first portion is coupled to a second portion. The second portion has a second internal face with a convex portion. A plurality of adjustable pins is positioned along the internal faces of the first portion and the second portion. The pins have a height that is adjustable with respect to the internal faces. A locking mechanism is included for securing the first portion to the second portion. The system includes one or more air inlets and one or more air exits and a source of cooling air coupled to the one or more air inlets.

Abstract: A turbine nozzle having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, and within an envelope of approximately +/?0.049 inches, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form the airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.

Abstract: A turbine blade having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form an airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.

Abstract: A turbine blade having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form an airfoil shape. The X, Y, and Z distances may be scaled as a function of the same constant number and the X, Y, and Z distances lie within an envelope of approximately +/?0.032 inches in directions normal to the airfoil.