Abstract: A labyrinth seal that may include a first leg positioned about a high-pressure side of the dovetail tab, a second leg positioned about a low-pressure side of the dovetail tab, and a labyrinth chamber positioned between the first leg and the second leg. High-pressure fluid passing through the gap about the first leg expands within the labyrinth chamber so as to limit an amount of the high-pressure fluid that passes beyond the second leg.

Abstract: A method of fabricating a component for a gas turbine engine is provided. The method includes applying a bond coat to at least a portion of the component, applying a dense vertically cracked (DVC) thermal barrier coating to at least a portion of the bond coat using a spray mechanism positioned a first distance from the component, and overlying at least a portion of the DVC thermal barrier coating with a soft coat thermal barrier coating using a spray mechanism that is positioned a second distance away from the component, wherein the second distance is greater than the first distance to facilitate adherence of the soft coating thermal barrier coating to the DVC thermal barrier coating.

Abstract: A method of fabricating a component for a gas turbine engine is provided. The method includes applying a bond coat to at least a portion of the component, applying a dense vertically cracked (DVC) thermal barrier coating to at least a portion of the bond coat using a spray mechanism positioned a first distance from the component, and overlying at least a portion of the DVC thermal barrier coating with a soft coat thermal barrier coating using a spray mechanism that is positioned a second distance away from the component, wherein the second distance is greater than the first distance to facilitate adherence of the soft coating thermal barrier coating to the DVC thermal barrier coating.

Abstract: A labyrinth seal that may include a first leg positioned about a high-pressure side of the dovetail tab, a second leg positioned about a low-pressure side of the dovetail tab, and a labyrinth chamber positioned between the first leg and the second leg. High-pressure fluid passing through the gap about the first leg expands within the labyrinth chamber so as to limit an amount of the high-pressure fluid that passes beyond the second leg.

Abstract: A pattern for improving aerodynamic performance of a turbine includes a material disposed in a pattern at a base surface of a turbine shroud such that the material is capable of abradable contact with a tip portion of a turbine bucket. The pattern includes a first plurality of ridges disposed at the base surface such that a first portion of the first plurality of ridges corresponding to a back portion of the turbine bucket is oriented at a first angle with respect to an axis of rotation of the turbine bucket. Each ridge of the first plurality of ridges has a first sidewall and a second sidewall having a first end and a second end. The first ends of the first and second sidewalls extend from the base surface. The first and second sidewalls slope toward each other with substantially equal but opposite slopes until meeting at the second ends of respective first and second sidewalls defining a centerline and a top portion of the ridge.

Abstract: An article of manufacture pattern for improving aerodynamic performance of a turbine including an abradable material capable of abradable contact. The abradable material is disposed in a pattern. The pattern includes a first plurality of ridges disposed at a base surface of the turbine. Each ridge of the first plurality of ridges has a first sidewall and a second sidewall having a first end and a second end. The first ends of the first and second sidewalls extend from the base surface. The first and second sidewalls slope toward each other with substantially equal but opposite slopes until meeting at the second ends of respective first and second sidewalls defining a centerline and a top portion of the ridge.

Abstract: An abradable seal is provided to improve turbine performance by physically reducing the clearance between a flange portion of the nozzle and an opposed angel wing/seal plate member of the bucket. The provision of an abradable seal also mitigates angel wing/seal plate tooth or fin wear by providing for abradable contact without metal to metal hard rub.

Abstract: A coating and method for overcoating a TBC on a component used in a high-temperature environment, such as the combustor section of an industrial gas turbine. The coating defines the outermost surface of the component and is formed of at least two layers having different compositions. An inner layer of the coating contains alumina in a first silica-containing matrix material that is free of zinc titanate. An outer layer of the coating contains alumina, a glass material, and zinc titanate in a second silica-containing matrix material. The outer layer of the coating has a surface roughness of not greater than three micrometers Ra and forms the outermost surface of the component. The coating reduces the component temperature by reducing the convective and radiant heat transfer thereto.

Abstract: An abradable seal is provided to improve turbine performance by physically reducing the clearance between a flange portion of the nozzle and an opposed angel wing/seal plate member of the bucket. The provision of an abradable seal also mitigates angel wing/seal plate tooth or fin wear by providing for abradable contact without metal to metal hard rub.