An cicctrodcposited nickel-chromium-aluminum (NiCrAl) 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 NiCrAl composite is heat-treated to form an aluminum compound and to restore materials lost during repair processes of the turbine components.

A nickel-chromium (NiCr) alloy and a method for electrodepositing the NiCr 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 NiCr alloy including from 2 to 50 wt % chromium balanced with nickel. The turbine component coated with the NiCr 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.

An objective of the invention to provide an alloy article that exhibits even better mechanical properties and/or even higher corrosion resistance than conventional high entropy articles without sacrificing the attractive properties thereof, a product formed of the alloy article, and a fluid machine having the product. An alloy article according to the invention has a predetermined chemical composition comprising: Co, Cr, Fe, Ni and Ti, each within a range of 5 atomic % or more and 35 atomic % or less; Mo within a range of more than 0 atomic % and less than 8 atomic %; an element with a larger atomic radius than the atomic radiuses of Co, Cr, Fe and Ni within a range of more than 0 atomic % and 4 atomic % or less; and a balance of inevitable impurities.

A gas turbine engine includes: a turbine section including a casing extending circumferentially about a plurality of turbine blades and having at least one seal member coated with an abradable coating. At least one turbine blade has sides and a tip and at least one seal member is located adjacent to the tip of the at least one turbine blade. The tip of the at least one turbine blade has a wear resistant layer and an abrasive coating disposed on the wear resistant layer. The wear resistant layer has a thickness less than or equal to 10 mils (254 micrometers) and includes metal boride compounds.

A gas turbine engine includes an engine static structure extending circumferentially about an engine centerline axis; a compressor section, a combustor section, and a turbine section within the engine static structure. At least one of the compressor section and the turbine section includes at least one airfoil and at least one seal member adjacent to the at least one airfoil. A tip of the at least one airfoil is metal having a wear resistant coating and the at least one seal member is coated with an abradable coating. The wear resistant coating is formed as a layer in a base metal surface of the airfoil, has a thickness less than or equal to 10 mils (254 micrometers) and includes metal boride compounds.

A heat-resistant cast steel contains 0.55 mass % or more and 1.0 mass % or less C, 1.5 mass % or more and 3.5 mass % or less Si, more than 0 mass % and 2 mass % or less Mn, 6 mass % or more and 11 mass % or less Ni, 22 mass % or more and 27 mass % or less Cr, and more than 0 mass % and 0.6 mass % or less No, and the balance being Fe and unavoidable impurities.

A gas turbine that includes a rotor blade that includes an airfoil. The airfoil may include non-integral portions in which: a base portion is made from a first material; and a top portion is made from a second material. The airfoil may include a connector securing the base portion to the top portion of the airfoil, wherein the connector comprises a wire-lock connector that includes: a tab extending from one of the base portion and the top portion; a complimentary slot for receiving the tab formed in the other one of the base portion and the top portion; a first groove formed in a side of the tab; a second groove formed in a side of the slot; a retaining aperture formed cooperatively via an alignment of the first and second grooves upon the tab being received into the slot; and a retaining wire housed within the retaining aperture.

A seal for sealing a radially outer component of a gas turbine engine stator to a radially inner component thereof includes an axially resilient seal carrier adapted for mounting the seal carrier and including at a radially inner portion thereof, a pair of radially spaced, axially extending, radially resilient jaws adapted to clamp a sealing element such as a rope seal there between in sealing engagement with a radially inner component of the engine stator.

A rocket engine has a combustion chamber having an inlet and an outlet, the inlet fluidly connectable to a source of oxidizer, the outlet in fluid communication with an environment outside the combustion chamber for expelling combustion gases, a first fuel having a first solid propellant and a second fuel having a second solid propellant, the first and second fuels located within the combustion chamber and configured to be exposed to the oxidizer injected in the combustion chamber via the inlet, the first solid propellant having a regression rate greater than that of the second solid propellant.

A hybrid rocket engine system has: an oxidizer tank containing a liquid oxidizer; a rocket engine having a combustion chamber operatively connected to the oxidizer tank; a solid propellant fuel within the combustion chamber; a nozzle fluidly connected to the combustion chamber, the nozzle having a convergent section and a divergent section downstream of the convergent section; and a thrust vector control device operatively connected to the divergent section of the nozzle and operable to inject a fluid through at least one aperture defined through the divergent section for controlling a direction of a thrust generated by the rocket engine.