A method and casting core for forming a landing for welding a baffle inserted into an airfoil are disclosed, wherein the baffle landing of the blade or vane is formed in investment casting by the casting core rather than by wax, reducing tolerances and variability in the location of the baffle inserted into the cooling cavity of airfoil when the baffle is welded to the baffle landing.

A gas turbine engine system includes a compressor, gas turbine, and combustor including a plurality of late lean fuel injectors supplied with secondary fuel to its interior. The gas turbine engine system includes a wash system in communication with the late lean fuel injectors. The wash system includes a water source; water pump; anti-corrosion agent fluid source with an anti-corrosion agent including a polyamine corrosion inhibitor; anti-corrosion agent supply piping in fluid communication with the anti-corrosion agent fluid source; mixing chamber receiving water and anti-corrosion agent to produce an anti-corrosion mixture in fluid communication with the mixing chamber and the plurality of late lean fuel injectors. Fluid from the mixing chamber including the water, the anti-corrosion agent fluid source, or a mixture thereof is injected, while the gas turbine engine is off-line, into the combustor at at least one of the plurality of late lean fuel injectors.

A gas turbine engine system includes a gas turbine engine including a compressor, combustor including a plurality of late lean fuel injectors supplied with secondary fuel; gas turbine, and wash system configured to be attached and in fluid communication with the late lean fuel injectors. The wash system includes a water source including water; first fluid source including a first fluid providing vanadium ash and vanadium deposit mitigation and removal from internal gas turbine components; a mixing chamber in communication with the water source and first fluid source; a water pump to pump the water to the mixing chamber; a first fluid pump the first fluid to the mixing chamber; a fluid line in fluid communication with the mixing chamber and late lean fuel injectors so fluid from the mixing chamber is injected into the combustor at the late lean fuel injectors while the gas turbine engine is on-line.

A system comprises a gas turbine engine including a compressor, combustor, gas turbine, the combustor including a plurality of late lean fuel injectors; and wash system configured to be attached to and in fluid communication with the a plurality of late lean fuel injectors of the combustor. The wash system includes a water source supplying water; a first fluid source supplying a first fluid; a mixing chamber in communication with the water source and first fluid source; a water pump to pump water to the mixing chamber; a first fluid pump to pump the first fluid to the mixing chamber; a fluid line in fluid communication with the mixing chamber and at least one of the plurality of late lean fuel injectors so fluid from the mixing chamber is injected into the combustor at late lean fuel injectors. The wash system is operated with the gas turbine engine off-line.

A component according to the disclosure may include a body having an aperture therein for receiving one of a turbomachine shaft or a lathe chuck, wherein in response to the body being coupled to the lathe chuck, the aperture is oriented substantially axially relative to an axis of rotation of the body with the lathe chuck; and a flange coupled to and in direct axial contact with the body, the flange including a surface that extends axially relative to the axis of rotation of the body, wherein the surface of the flange comprises a matingly engageable face configured to contact an axially aligned surface during operation of the component and having a sanding indentation thereon, wherein a surface roughness of the surface of the flange is less than a surface roughness of a remainder of the component.

A gas turbine engine system includes a compressor, gas turbine, and combustor including a plurality of late lean fuel injectors supplied with secondary fuel to its interior. The gas turbine engine system includes a wash system in communication with the late lean fuel injectors. The wash system includes a water source; water pump; anti-corrosion agent fluid source with an anti-corrosion agent including a amine corrosion inhibitor; anti-corrosion agent supply piping in fluid communication with the anti-corrosion agent fluid source; mixing chamber receiving water and anti-corrosion agent to produce an anti-corrosion mixture in fluid communication with the mixing chamber and the plurality of late lean fuel injectors. Fluid from the mixing chamber including the water, the anti-corrosion agent fluid source, or a mixture thereof is injected, while the gas turbine engine is off-line, into the combustor and at least one of the plurality of late lean fuel injectors.

The invention relates to a nickel-based superalloy comprising, in percentages by mass, 4.0 to 5.5% rhenium, 1.0 to 3.0 ruthenium, 2.0 to 14.0% cobalt, 0.3 to 1.0% molybdenum, 3.0 to 5.0% chromium, 2.5 to 4.0% tungsten, 4.5 to 6.5% aluminum, 0.50 to 1.50% titanium, 8.0 to 9.0% tantalum, 0.15 to 0.30% hafnium, 0.05 to 0.15% silicon, the balance being nickel and unavoidable impurities.

The invention also relates to a single-crystal blade (20A, 20B) comprising such an alloy and a turbomachine (10) comprising such a blade (20A, 20B).

A method for manufacturing a part coated with a protective coating includes: forming a protective coating across all or part of the surface of a part, wherein the part includes a refractory alloy including a niobium matrix containing metal silicide inclusions, wherein the protective coating is formed by a pack carburization method from a cement including: i. a mixture A of (Nb.sub.xTi.sub.1-x).sub.3M.sub.3CrSi.sub.6 and M.sub.0.6Cr.sub.0.4Si where M denotes Fe, Co or Ni and x is between 0 and 1, or ii. a mixture B of M′Si, NbSi.sub.2 and Nb.sub.4M′.sub.4Si.sub.7 where M′ denotes Fe, Co or Ni.

Core assemblies and methods for manufacturing components of gas turbine engines include a first core body having a first trunk configured to attach to a first location of a cavity core structure, a first branch extending from the first trunk and configured to form a first portion of a first cooling circuit, the first branch having a first joining surface, and a second core body having a second trunk configured to attach to a second location of a cavity core structure, a first branch of the second core body extending from the second trunk and configured to form a first portion of a second cooling circuit in the component. The first branches of the core bodies joined to form a junction. The junction defines a merger of the first cooling circuit and the second cooling circuit proximate to an exit of the first and second cooling circuits.

Airfoil assemblies for gas turbine engines are described. The airfoil assemblies include an airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side, the airfoil body extending in a radial direction between a first end and a second end, wherein the airfoil defines an internal cavity bounded by interior surfaces of the airfoil body, the airfoil body formed from a high-temperature-material material and a metallic insert member installed within the internal cavity. One or more radially extending ribs are arranged on an exterior surface of the metallic insert member and defining one or more radially extending passages between the exterior surface of the metallic insert member and the interior surface of the airfoil body.