United Technologies Corporation

David R Malley

Gary M Lomasney

David A Rutz

William M Rose

Monika D Kinstler

James J Moor

Daniel F Paulonis

Robert P Schaefer

Edgar E Brown

David W Anderson

William H King

Carl E Kelly

John F Mullooly

Steven Ivory

Jose A Echevarria

Kevin C McMahon

Joseph Rokicki

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.

Compressor blade flow form technique for repair

A method for repairing a turbine engine component comprises the steps of providing a turbine engine component formed from a nickel-based superalloy, removing any defects from the turbine engine component, the removing step comprising subjecting the turbine engine component to a solutioning heat treatment using a heating rate in the range of from 29 degrees Fahrenheit/minute to approximately 40 degrees Fahrenheit/minute as the component passes through a temperature in the range of from 1100 degrees Fahrenheit to 1600 degrees Fahrenheit, and welding the turbine engine component to effect the repair. A deliberately slow post-solution treatment cooling rate is used to substantially prevent weld cracks.

Enhanced weldability for high strength cast and wrought nickel superalloys

For a group of nickel-based superalloys, improved properties have been obtained by stabilizing at increased temperature for a reduced time relative to prior art specifications. In particular, improved creep properties have been obtained with a one-hour 1800&#xb0; F. stabilization relative to a prior art four-hour 1500&#xb0; F. stabilization.

Superalloy stabilization

A method of repairing a braze joint and the resulting joint includes removing an unsuitable portion of a first joining material in a joint between a first member and a second member of an aerospace assembly and cleaning the joint. A stripping solution is used to remove the unsuitable portion of the first joining material. A first cleaning solution is used to clean the joint, and the joint is further cleaned using a second cleaning solution. During a rebrazing step, a second joining material moves into a void left by the removed unsuitable portion from the removal step to form a new joint between the first member and the second member that includes the remaining first joining material and the second joining material.

Repair of braze joint and article repaired

Systems and methods for repairing Thermo-Span&#xae; gas turbine engine components are described herein. Embodiments of these methods minimize post-weld residual stresses in a weld repaired Thermo-Span&#xae; component by solution heat treating the component by heating the component to about 2000&#xb0; F.&#xb1;25&#xb0; F., holding the component at about 2000&#xb0; F.&#xb1;25&#xb0; F. for about one hour; and cooling the component to below about 700&#xb0; F. at a rate equivalent to cooling in air; and precipitation heat treating the component by heating the component to about 1325&#xb0; F.&#xb1;25&#xb0; F., holding the component at about 1325&#xb0; F.&#xb1;25&#xb0; F. for about 8 hours, cooling the component to about 1150&#xb0; F.&#xb1;25&#xb0; F. at a maximum rate of about 100&#xb0; F./hour, holding the component at about 1150&#xb0; F.&#xb1;25&#xb0; F. for about 8 hours, and cooling the component at a predetermined cooling rate. Dimensions of the fully-machined and weld repaired component are maintained during solution heat treating and precipitation heat treating via custom designed furnace tools.

Methods for repairing gas turbine engine components

A high strength nickel base superalloy article having a machined surface is disclosed. The superalloy comprises, in weight percent, 1.2-3.5 Al, 3.0-7.0 Ti, 12.0-20.0 Cr, 2.0-3.9 Mo, 10.0-20.0 Co, 0-4.5 W, 0.005-0.025 C, 0.005-0.05 B, 0.01-0.1 Zr, 0-0.005 Mg, 0-1.0 Ta, 0-1.0 Nb, 0-2.0 Fe, 0-0.3 Hf, 0-0.02 Y, 0-0.1 V, 0-1.0 Re, balance essentially Ni. The superalloy further comprises a plurality of discrete carbides essentially free from molybdenum for increased fatigue strength.

Low cycle fatigue strength nickel base superalloys

Quench cracking of nickel base superalloy articles is avoided by plating the articles with a layer of nickel or similar metal prior to heat treatment.

Elimination of quench cracking in superalloy disks

Forging processes are described for cast superalloys. The forgeability of superalloy materials is improved by the development of overaged microstructure. The total forging process is usually performed in a series of steps with intervening heat treatments to reform the overaged microstructure.

Forging process for superalloys

A fushion welded surfacing process for use with crack prone substrates is described. The process involves the application of an intermediate layer of crack resistant alloy by fusion welding followed by the application of the desired final surfacing alloy by fusion welding. The intermediate layer is an age hardenable nickel superalloy to which intentional additions of manganese have been made. The final surfacing alloy may be of the hardfacing type or may be of a type which imparts corrosion and oxidation resistance to the underlying article.

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Bolton, CT, United States of America

David R Malley