A process for manufacturing a turbomachine part coated with a thermal barrier, includes manufacturing the part by additive manufacture; electrophoretic depositing the part of a layer including particles of a ceramic material; consolidating the layer by heat treatment to obtain a ceramic coating.

An article of manufacture includes an article body portion (36); at least one pedestal (58A, 60A, 84, 86) integrally formed with the article body portion (36), and disposed at an outer periphery of (52), and structurally coupled to the article body portion (36); and at least one internal feature (72) disposed at least partially within the article body portion (36) and at least partially within the pedestal (58A, 60A, 84, 86). At least a portion of the internal feature (72) is hollow.

A method for manufacturing a metal component includes the following steps: a shell made of a titanium-based material is provided, the shell having a first surface and a second surface remote from the first surface; a covering layer made of a titanium fire-resistant material is produced by additive manufacturing on the shell such that the covering layer at least partially covers the first surface and/or the second surface; and, after the additive manufacturing step, the metal component is heat treated at a temperature of between 200° C. and 1000° C.

A production method for an alloy member having mainly high hardness and high resistance to corrosion and produced by an additive manufacturing method, the alloy member, and a product using the alloy member are provided. The production method for an alloy member includes: an additive manufacturing step of forming a shaped member through an additive manufacturing method using an alloy powder containing elements Co, Cr, Fe, Ni, and Ti each in a range of 5 atom% to 35 atom% and containing Mo in a range exceeding 0 atom% and 8 atom% or less, the remainder being unavoidable impurities; and a heat treatment step of holding the shaped member in a temperature range higher than 500° C. and lower than 900° C. directly after the additive manufacturing step without undergoing a step of holding the shaped member in a temperature range of 1080° C. to 1180° C.

An airfoil component for attaching to a cropped airfoil is provided. The cropped airfoil comprises a cropped airfoil attachment section and a cropped first side opposite a cropped second side, which each extend axially between a cropped first edge and a cropped second edge to define a cropped chord length. The airfoil component comprises a body having a component first side opposite a component second side. The body defines an attachment section for attaching the airfoil component to the cropped airfoil at the cropped airfoil attachment section. The attachment section extends axially between a component first edge and a component second edge to define a component chord length, and the attachment section is oversized with respect to the cropped airfoil attachment section such that the component chord length is longer than the cropped chord length. Systems and methods also are provided.

Thermal insulation blankets and methods of using the same to repair articles, such as components of turbine engines are disclosed. The thermal insulation blankets are a multilayer composite material that includes: a first flexible fiberglass fabric layer having two sides that are coated with a coating that covers the fiberglass fibers; an intermediate second fiberglass insulation blanket layer made of continuous glass fibers; and a third layer of flexible heat reflective material. The method involves wrapping a component, such as a component of a turbine engine, with one or more thermal insulation blankets in order to retain heat within the confines of a prescribed area for installing or removing parts with a precision fitment. The blankets allow enough heat to induce even thermal expansion of the components.

A blade for an aircraft gas turbine engine includes, in a longitudinal direction, a blade root, a shank and an aerofoil body, the aerofoil body extending in the longitudinal direction between the shank and a blade tip and in a transverse direction between a leading edge made of metal material and a trailing edge. The blade includes a blade core made of composite material having a three-dimensional woven fibrous reinforcement forming the blade root, the shank and a part of the aerofoil body. The blade also includes a skin made of composite material having a two-dimensional woven fibrous reinforcement surrounding the aerofoil body part of the blade core, the skin being interposed between the leading edge made of metal material and a front edge of the aerofoil body part of the blade core to define a thinned leading edge portion, the skin including one or more two-dimensional woven plies.

An assembly for hot forming a component includes first and second location features reversibly attached to a sheet metal element that forms the component. The assembly is loaded into a high temperature forming rig by loosely locating the location features into corresponding datum features in a forming tool. As the assembly heats up, the location features locate accurately with the datum features. Thus, the assembly can be accurately located onto the tool accurately and with minimal opening of the hot forming rig, thereby improving efficiency and safety.

A method of localized repair to a damaged thermal barrier, the method including subjecting a part coated in a damaged thermal barrier to electrophoresis treatment, the part being made of an electrically conductive material, the damaged thermal barrier including a ceramic material and presenting at least one damaged zone that is to be repaired, the part being present in an electrolyte including a suspension of particles in a liquid medium, the ceramic coating being deposited by electrophoresis in the damaged zone in order to obtain a repaired thermal barrier for use at temperatures higher than or equal to 1000° C., the particles being made of a material different from the ceramic material present in the damaged thermal barrier.

According to an embodiment, a method of correcting a bend of a joint type-turbine rotor comprises: measuring displacement of a convex portion of the bend at a joined portion of the joint type-turbine rotor or displacement of a surface opposite to the convex portion in a circumferential direction of the joint type-turbine rotor; heating the convex portion; and cooling the joined portion after the step of heating. The steps of heating and cooling are performed during the step of measuring.