Abstract: A component for a turbine engine comprises a wall with a surface along which a hot airflow passes, a second surface along which a cooling airflow passes, and an insert mounted to the wall wherein the material used for the insert can have a higher temperature capability than that of the wall.

Abstract: An assembly for a turbine engine comprises a circumferential band defining an air flow path, a component having a leading edge and extending from the band into the flow path, and a trough located in a surface of the band along the leading edge.

Abstract: Flow path assemblies and methods for assembling such assemblies are provided. For example, a flow path assembly comprises a turbine nozzle segment including an airfoil extending axially between leading and trailing edges, an inner band defining a portion of a flow path inner boundary, and an outer band defining a portion of a flow path outer boundary. The airfoil includes a trailing edge portion defining the trailing edge and extending axially beyond the inner and outer bands such that the trailing edge is defined axially aft of the inner and outer band aft ends. The flow path assembly further comprises a shroud having a forward portion that extends axially along the trailing edge portion such that the forward portion defines the flow path outer boundary at the trailing edge portion. Methods of assembling flow path assemblies having a turbine nozzle assembly and inner and outer members also are provided.

Abstract: A sealing system for sealing a gap between two adjacent components includes a first component including a sealing face, a second component, and a seal cap. The second component includes a seal ridge extending from a surface of the second component towards the sealing face, and a fluid conduit extending through the second component and the seal ridge, the fluid conduit configured to channel a seal activating fluid from a fluid source. The seal cap is configured to matingly engage the seal ridge, and includes an end wall positionable between the sealing face and the seal ridge, and a pair of seal legs extending from the end wall towards the surface of the second component.

Abstract: Turbine nozzle segments and systems are provided. For example, a turbine nozzle segment comprises an airfoil extending axially between leading and trailing edges; an inner band defining a portion of a flow path inner boundary; and an outer band defining a portion of a flow path outer boundary. An airfoil trailing edge portion defines the trailing edge and extends axially beyond the inner and outer bands such that the trailing edge is defined axially aft of the inner and outer band aft ends. An exemplary turbine nozzle system comprises the turbine nozzle segment, an outer member positioned aft of the outer band at the trailing edge portion outer end to define the flow path outer boundary along the trailing edge portion, and an inner member positioned aft of the inner band at the trailing edge portion inner end to define the flow path inner boundary along the trailing edge portion.

Abstract: A method of forming a structure in a turbine component having a seal slot, the slot including walls defining an opening therebetween, the method includes the step of using an additive manufacturing process to form a neck structure on a wall so as to reduce a size of the opening.

Abstract: A compliant seal component, a turbomachine including such compliant seal component, and an associated method for cooling the compliant seal component disposed in the turbomachine are disclosed. The compliant seal component includes a convoluted portion and a plurality of end portions spaced apart from each other. The plurality of end portions is joined to the convoluted portion. An end portion of the plurality of end portions includes a first section and a second section, the second section defining a plenum. The end portion further includes a plurality of inlet holes and a plurality of outlet holes. The plurality of inlet holes is configured to direct a portion of a cooling fluid to the plenum. The portion of the cooling fluid in the plenum is configured to cool the end portion. The plurality of outlet holes is configured to discharge the portion of the cooling fluid from the plenum.

Abstract: An airfoil for a gas turbine engine includes a pressure side and a suction side, with a root and a tip wall. The pressure side and suction side extend beyond the tip wall to define a tip channel, defining a plurality of internal and external corners. A tip shelf can be defined in the pressure or suction sidewalls. A film hole can extend to the tip shelf, such that the film hole can be shaped to direct a flow of cooling fluid to the film hole to improve film cooling at the tip shelf.

Abstract: A gas turbine engine arcuate segment includes arcuate flange with anti-chording means extending away from annular wall. Anti-chording means may include insert in or bonded to flange and made of different alpha material than the annular wall. Anti-chording means may be heating means for heating flange. Heating means includes hot air inlet to and an outlet from a circumferentially extending heating flow passage embedded in flange and may further include a cold air inlet to heating flow passage. Heating flow passage may be a serpentine heating flow passage with an undulating heating flowpath. A turbine nozzle segment includes one or more airfoils extending radially between inner and outer arcuate band segments and forward and aft outer flanges extending radially from the outer arcuate band segment include anti-chording means. An arcuate turbine shroud segment includes forward and aft shroud rail segments with anti-chording means.

Abstract: A compliant seal component, a turbomachine including such compliant seal component, and an associated method for cooling the compliant seal component disposed in the turbomachine are disclosed. The compliant seal component includes a convoluted portion and a plurality of end portions spaced apart from each other. The plurality of end portions is joined to the convoluted portion. An end portion of the plurality of end portions includes a first section and a second section, the second section defining a plenum. The end portion further includes a plurality of inlet holes and a plurality of outlet holes. The plurality of inlet holes is configured to direct a portion of a cooling fluid to the plenum. The portion of the cooling fluid in the plenum is configured to cool the end portion. The plurality of outlet holes is configured to discharge the portion of the cooling fluid from the plenum.

Abstract: Turbine nozzle segments and systems are provided. For example, a turbine nozzle segment comprises an airfoil extending axially between leading and trailing edges; an inner band defining a portion of a flow path inner boundary; and an outer band defining a portion of a flow path outer boundary. An airfoil trailing edge portion defines the trailing edge and extends axially beyond the inner and outer bands such that the trailing edge is defined axially aft of the inner and outer band aft ends. An exemplary turbine nozzle system comprises the turbine nozzle segment, an outer member positioned aft of the outer band at the trailing edge portion outer end to define the flow path outer boundary along the trailing edge portion, and an inner member positioned aft of the inner band at the trailing edge portion inner end to define the flow path inner boundary along the trailing edge portion.

Abstract: A compliant seal component, a turbomachine including such compliant seal component, and an associated method for cooling the compliant seal component disposed in the turbomachine are disclosed. The compliant seal component includes a convoluted portion and a plurality of end portions spaced apart from each other. The plurality of end portions is joined to the convoluted portion. An end portion of the plurality of end portions includes a first section and a second section, the second section defining a plenum. The end portion further includes a plurality of inlet holes and a plurality of outlet holes. The plurality of inlet holes is configured to direct a portion of a cooling fluid to the plenum. The portion of the cooling fluid in the plenum is configured to cool the end portion. The plurality of outlet holes is configured to discharge the portion of the cooling fluid from the plenum.

Abstract: A blade for a gas turbine engine comprises an airfoil having a pressure side and a suction side, with a root and a tip wall. The pressure side and suction side extend beyond the tip wall to define a tip channel, defining a plurality of internal and external corners. The corners comprise fillets to define a thickness being greater than the thickness for the pressure, suction, or tip walls. A film hole can extend through the fillet, such that the length of the film hole at the fillet can be increased to define an increased length-to-diameter ratio for the film hole to improve film cooling through the film hole.

Abstract: An inertial separator for removing particles from a fluid stream includes a body defining a through passage, with at least one separator inlet and at least one separator outlet, a turn provided in the body between the at least one separator inlet and the at least one separator outlet, and a flow splitter having at least one particle outlet fluidly coupled to the outside of the through passage.

Abstract: An airfoil for a gas turbine engine can have an exterior wall and an interior wall, with each wall having a thickness. The walls can intersect to define a corner at the intersection. A cooling passage can be defined by the walls at or near the corner to provide fluid communication between the interior and exterior of the airfoil. A film hole can be disposed in the walls and can have a length and diameter to define a ratio of length to diameter, L/D. An arcuate fillet can be located in the corner to define an effective radius for the fillet. The effective radius can be at least 1.5 times larger than the thicknesses of the walls to provide for an increased length to diameter ratio for the film hole.

Abstract: A centrifugal compressor apparatus includes: an impeller mounted for rotation about a central longitudinal axis, the impeller including an impeller disk carrying an array of impeller blades around its periphery; a diffuser disposed downstream of the impeller, configured to diffuse and turn airflow discharged from the impeller; an extraction scoop disposed in fluid communication with the impeller and with a plenum disposed adjacent the impeller, and a dirt collector disposed in fluid communication with the plenum and configured to trap dirt therein.

Abstract: A filter system and method use a filter housing that defines an interior chamber and that includes an inlet opening extending into the interior chamber. The outer air flow housing has an outlet conduit through which a flow of air having particles is directed toward the inlet opening of the filter housing along a flow direction toward the interior chamber of the filter housing. The outer air flow housing engages the filter housing such that the filter housing is separated from the outer air flow housing along the flow direction to permit at least some of the air to pass around an exterior of the filter housing and exit the outer air flow housing while the particles in the at least some of the air pass into the interior chamber of the filter housing through the inlet opening.

Abstract: A blade for a gas turbine engine comprises an airfoil having a pressure side and a suction side, with a root and a tip wall. The pressure side and suction side extend beyond the tip wall to define a tip channel, defining a plurality of internal and external corners. The corners comprise fillets to define a thickness being greater than the thickness for the pressure, suction, or tip walls. A film hole can extend through the fillet, such that the length of the film hole at the fillet can be increased to define an increased length-to-diameter ratio for the film hole to improve film cooling through the film hole.

Abstract: An airfoil comprises at least one wall defining a leading edge, a trailing edge, a pressure side extending between the leading edge and the trailing edge, and a suction side extending between the leading edge and the trailing edge. The airfoil is curved in three dimensions and has one or more cavities defined by an interior surface of the at least one wall. A plurality of cooling film holes extending between the cavity and at least one cooling trench located on at least one of the pressure side and the suction side, spaced from the leading edge. The at least one trench has a floor spaced from an outer surface of the airfoil. The plurality of cooling film holes extend through the floor at an angle other than perpendicular.

Abstract: A turbomachinery component with a surface that includes a bounded wear coat, the component includes: a body; a contact surface defined by the body; a recess extending into the body and communicating with the contact surface; and a wear coat positioned in the recess.