Abstract: A nickel-chromium (Ni—Cr) alloy and a method for electrodepositing the Ni—Cr alloy on a turbine engine component for dimensionally restoring the engine component are described. The engine component is restored by rebuilding wall thickness with the Ni—Cr alloy including from 2 to 50 wt % chromium balanced with nickel. The turbine component coated with the Ni—Cr alloy is heat-treated at a high temperature to homogenize composition of the alloy to mimic the base alloy and to restore materials lost during repair of the turbine component.

Abstract: An cicctrodcposited nickel-chromium-aluminum (Ni—Cr—Al) composite including nickel- chromium alloy and aluminum, and alloys or compounds formed by Al, Cr and Ni applied on turbine components comprises from 2 to 50 wt % chromium, from 0.1 to 6 wt % aluminum, and a remaining balance of nickel, wherein the Ni—Cr—Al composite is heat-treated to form an aluminum compound and to restore materials lost during repair processes of the turbine components.

Abstract: A nickel-chromium (Ni—Cr) alloy and a method for electrodepositing the Ni—Cr alloy on a turbine engine component for dimensionally restoring the engine component are described. The engine component is restored by re-building wall thickness with the Ni—Cr alloy including from 2 to 50 wt % chromium balanced with nickel. The turbine component coated with the Ni—Cr alloy is heat-treated at a high temperature to homogenize composition of the alloy to mimic the base alloy and to restore materials lost during repair of the turbine component.

Abstract: An electrodeposited nickel-chromium-aluminum (Ni—Cr—Al) composite including nickel-chromium alloy and aluminum, and alloys or compounds formed by Al, Cr and Ni applied on turbine components comprises from 2 to 50 wt % chromium, from 0.1 to 6 wt % aluminum, and a remaining balance of nickel, wherein the Ni—Cr—Al composite is heat-treated to form an aluminum compound and to restore materials lost during repair processes of the turbine components.

Abstract: Aspects of the disclosure are directed to repairing or restoring a liner associated with a combustor of an aircraft engine. The liner is at least partially stripped of at least one detail. An automated welding operation is applied to the liner. A coating of the liner on a cold side of the liner is retained on the liner during the welding operation.

Abstract: An electrodeposited nickel-chromium-aluminum (Ni—Cr—Al) composite including nickel-chromium alloy and aluminum, and alloys or compounds formed by Al, Cr and Ni applied on turbine components comprises from 2 to 50 wt % chromium, from 0.1 to 6 wt % aluminum, and a remaining balance of nickel, wherein the Ni—Cr—Al composite is heat-treated to form an aluminum compound and to restore materials lost during repair processes of the turbine components.

Abstract: A nickel-chromium (Ni—Cr) alloy and a method for electrodepositing the Ni—Cr alloy on a turbine engine component for dimensionally restoring the engine component are described. The engine component is restored by re-building wall thickness with the Ni—Cr alloy including from 2 to 50 wt % chromium balanced with nickel. The turbine component coated with the Ni—Cr alloy is heat-treated at a high temperature to homogenize composition of the alloy to mimic the base alloy and to restore materials lost during repair of the turbine component.

Abstract: A method of additively repairing a combustor liner panel includes removing a combustor liner panel from a combustor, inspecting the combustor liner panel to identify a damaged portion, removing material from the combustor liner panel around the damaged portion to form a repair zone having a substantially flat platform, and adding repair material to the repair zone on a layer by layer basis using an additive repair process.

Abstract: Aspects of the disclosure are directed to repairing or restoring a liner associated with a combustor of an aircraft engine. The liner is at least partially stripped of at least one detail. An automated welding operation is applied to the liner. A coating of the liner on a cold side of the liner is retained on the liner during the welding operation.

Abstract: A method of hole defect repair includes removing one or more defects at or near a desired hole in a part by removing a portion of the part proximate the desired hole in a series of chain link patterns, and welding a filler material to the part after removing the chain link pattern portion of the part.

Abstract: A method for working or repairing a combustor wall panel that may be damaged is disclosed. The method comprises providing a supplemental body to a combustor float wall panel, and attaching at least one cooling pin to the supplemental body.

Abstract: A liner panel for a gas turbine engine includes a cold side having a backside feature and a hot side with a variable polarity weld.

Abstract: A method of treating a surface includes the steps of providing a plating including at least some nickel over a nickel alloy surface in a thickness less than 0.0005? (0.001 cm), and exposing the surface to a fluoride ion cleaning to remove impurities on the surface, and leaving at least some of the plating.

Abstract: Methods for repairing gas turbine components damaged by corrosion that not only restores the components back to their original dimensions to ensure proper engine operations, but also have the added benefit of mitigating future corrosion attacks thereby extending the useful life of the components.

Abstract: A method for working or repairing a combustor wall panel that may be damaged is disclosed. The method comprises providing a supplemental body to a combustor float wall panel, and attaching at least one cooling pin to the supplemental body.

Abstract: A method for depositing a first material on a substrate includes providing the substrate in a deposition chamber. A molten body is formed between the substrate and a source of the first material by melting one or more second materials. A flow of the first material is passed through the molten body and from the molten body to the substrate as a vapor flow. An essentially non-expending portion of the molten body comprises an alloy having a melting temperature below a melting temperature of the first material.

Abstract: A gas turbine engine blade has a platform. A root depends from the platform and an airfoil extends from the platform. The airfoil has leading and trailing edges and pressure and suction sides. The blade has a substrate having a surface. A compressive stress exists below a first region of the surface. The first region extends over a majority of a streamwise perimeter of the airfoil at a location at a spanwise distance from the tip or more than 50% of a tip-to-platform span. A coating is on the surface including at the location.

Abstract: A method for repairing a turbine engine component is provided. The method broadly comprises the steps of providing a turbine engine component having an airfoil portion, applying a coating to the airfoil portion, and restoring a tip portion, a chord, and surfaces of the airfoil portion to original dimensions in a single operation. A system for repairing the turbine engine component is also described.

Abstract: A method for restoring blade dovetails to their original shape comprising the steps of removing damaged material from a blade dovetail at a damage site, physically depositing a metal at the damage site in an amount sufficient to replace the removed damaged material, and machining the blade dovetail to the original shape.

Abstract: A method of repairing a component of a gas turbine engine that includes a metallic substrate, an existing aluminide coating, and a diffusion layer formed in the metallic substrate adjacent to the coating. The method includes removing at least a portion of the existing aluminide coating, removing material forming the diffusion layer, applying a new metallic layer to the metallic substrate, and applying a new aluminide coating over the new metallic layer to form a new diffusion layer in the new metallic layer. The new metallic layer is a substantially homogeneous material that is substantially similar in chemical composition to that of the metallic substrate, and the new metallic layer forms a structural layer having a thickness selected to provide a specified contour to the component.