Abstract: An airfoil assembly for a gas turbine engine and a method of transferring load from the ceramic matrix composite (CMC) airfoil assembly to a metallic vane assembly support member are provided. The airfoil assembly includes a forward end and an aft end with respect to an axial direction of the gas turbine engine. The airfoil assembly further includes a radially outer end component including a radially outwardly-facing end surface having a non-compression load-bearing feature extending radially outwardly from the outwardly-facing end surface and formed integrally with the outer end component, the feature configured to mate with a complementary feature formed in a radially inner surface of a first airfoil assembly support structure, the feature selectively positioned orthogonally to a force imparted into the airfoil assembly.

Abstract: A nozzle, a nozzle hanger, and a ceramic to metal attachment system are provided. The ceramic to metal attachment system includes the nozzle, a ceramic matrix composite, and the nozzle hanger, a metal. The attachment system also includes a clamping member adjacent a second surface of the nozzle and a mounting member of the nozzle. The attachment system includes a plurality of attachment members securing the nozzle, the clamping member, and the nozzle hanger together. A sealing member of the nozzle hanger seals off an airfoil of the nozzle from adjacent airflow.

Abstract: Methods for positioning neighboring nozzles of a gas turbine engine are provided. A method includes assembling a first nozzle assembly. The first nozzle assembly includes a first nozzle and a first nozzle support structure. The method further includes assembling a second nozzle assembly. The second nozzle assembly includes a second nozzle and a second nozzle support structure. The method further includes adjusting the first nozzle assembly and the second nozzle assembly such that an engineering dimension between the first nozzle and the second nozzle is within a predetermined engineering tolerance, and joining the first nozzle support structure and the second nozzle support structure together.

Abstract: Shrouds and shroud assemblies for gas turbine engines are provided. A shroud includes a shroud body. The shroud body includes a forward surface, a rear surface axially spaced from the forward surface, an inner surface extending between the forward surface and the rear surface, and an outer surface extending between the forward surface and the rear surface and radially spaced from the inner surface. The shroud further includes a flange extending from the shroud body, and a bore hole defined in the flange. The bore hole extends generally circumferentially through the flange between a first opening and a second opening.

Abstract: A nozzle assembly is provided which is, in part, formed of a low coefficient of thermal expansion material. The assembly includes a nozzle fairing formed of the low coefficient of thermal expansion material and includes a metallic strut extending radially through the nozzle fairing. Load is transferred from the nozzle fairing to a static structure in either of two ways: first, the strut may receive the load directly and/or second, load may be transferred from the nozzle fairing to at least one of the inner and outer support rings. Further, the nozzle fairing and strut may allow for internal airflow for cooling.

Abstract: A turbine nozzle includes an arcuate inner band having opposed flowpath and back sides, and an aft flange extending outward from the back side; an arcuate outer band having opposed flowpath and back sides; an airfoil-shaped turbine vane extending between the flowpath sides of the inner and outer bands, wherein the inner and outer bands and the vane comprise a ceramic low-ductility material; and a metallic collar surrounding the aft flange.

Abstract: A gas turbine engine assembly includes first and second annular members having different first and second thermal expansion coefficients connected together with dual arm V brackets. Brackets include first and second arms angularly spaced apart from a bracket centerline and extending axially away from bracket bases attached to a first one of the first and second annular members. Arms are attached to a second one of the first and second annular members. A turbine frame includes struts extending between outer and inner rings. An annular mixer and centerbody substantially made from a ceramic matrix composite materials is connected to and supported by the outer and inner rings with first and second sets respectively of the dual arm V brackets. Bracket bases of the first and second sets are attached to the outer and inner rings respectively. Arms of the first and second sets are attached to mixer and centerbody respectively.

Abstract: A component and method of cooling a component are provided. The component includes a leading edge, a trailing edge, at least one cavity between the leading edge and the trailing edge, at least one diffusion member adjacent to the cavity. The diffusion member includes an inlet adjacent to the cavity, a metering zone adjacent to the inlet, a diffusion zone adjacent to the metering zone, and an outlet adjacent the diffusion zone and adjacent the trailing edge. The diffusion member provides up to about 70% reduction in flow and uniform cooling of the trailing edge of the component.

Abstract: A nozzle, a nozzle hanger, and a ceramic to metal attachment system are provided. The ceramic to metal attachment system includes the nozzle, a ceramic matrix composite, and the nozzle hanger, a metal. The attachment system also includes a clamping member adjacent a second surface of the nozzle and a mounting member of the nozzle. The attachment system includes a plurality of attachment members securing the nozzle, the clamping member, and the nozzle hanger together. A sealing member of the nozzle hanger seals off an airfoil of the nozzle from adjacent airflow.

Abstract: A ceramic matrix composite (CMC) component and a method of attaching a static seal to a ceramic matrix composite component are provided. The CMC component includes a first end and a second end. A CMC metal interface member is attached to the second end. The CMC metal interface member is operable to join to a static seal in a gas turbine.

Abstract: A support frame includes a rail engaging a hook of a composite article having plys. The rail included a rail contact surface generally conformal to and contacting a hook contact surface of the hook. The rail and hook contact surfaces being substantially normal to the plys beneath the hook contact surface. One embodiment of the article is a composite nozzle segment including one or more airfoils extending radially between arcuate radially outer and inner band segments having plys curved about a centerline axis and the rail engages a hook of the radially inner band segment. The hook and the rail may be straight or arcuate. The frame includes a central body extending axially aftwardly from the rail to an aft frame flange. A frame pin may be mounted on the frame, engage the inner band segment or one of the airfoils, and may extend into a cavity in the airfoil.

Abstract: An aircraft gas turbine engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments. Each of the basesheet segments includes a panel body between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. Leading and trailing edge ribs supporting the overlapping flanges are clipped together with longitudinally spaced apart and transversely extending forward and aft slits in retainers.

Abstract: An aircraft engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments having panel bodies between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges include first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. A basesheet assembly includes a frame includes first and second side rails supporting the first and second basesheet side edges. An aircraft gas turbine engine exhaust nozzle includes divergent seals with the basesheet assemblies in sealing engagement against divergent flaps along the first and second side rails.

Abstract: A support frame includes a rail engaging a hook of a composite article having plys. The rail included a rail contact surface generally conformal to and contacting a hook contact surface of the hook. The rail and hook contact surfaces being substantially normal to the plys beneath the hook contact surface. One embodiment of the article is a composite nozzle segment including one or more airfoils extending radially between arcuate radially outer and inner band segments having plys curved about a centerline axis and the rail engages a hook of the radially inner band segment. The hook and the rail may be straight or arcuate. The frame includes a central body extending axially aftwardly from the rail to an aft frame flange. A frame pin may be mounted on the frame, engage the inner band segment or one of the airfoils, and may extend into a cavity in the airfoil.

Abstract: An aircraft engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments having panel bodies between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges include first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. A basesheet assembly includes a frame includes first and second side rails supporting the first and second basesheet side edges. An aircraft gas turbine engine exhaust nozzle includes divergent seals with the basesheet assemblies in sealing engagement against divergent flaps along the first and second side rails.

Abstract: An aircraft gas turbine engine exhaust nozzle basesheet includes longitudinally extending plurality of basesheet segments. Each of the basesheet segments includes a panel body between segment leading and trailing edges and slidable sealing joints with slidingly un-restrained center surfaces between adjacent segment leading and trailing edges. Segmented first and second basesheet side edges first and second segment side edges of the basesheet segments respectively. Slidably sealingly engaged overlapping flanges at the segment leading and trailing edges include tacked together transversely spaced apart first and second distal ends of the overlapping flanges. Leading and trailing edge ribs supporting the overlapping flanges are clipped together with longitudinally spaced apart and transversely extending forward and aft slits in retainers.

Abstract: A method facilitates assembling a flap system for a gas turbine engine exhaust nozzle including at least one backbone assembly. The method comprises providing a basesheet including a pair of circumferentially-spaced sides coupled together by an upstream side and a downstream side, forming at least one relief cut in the basesheet that extends at least partially across the basesheet from at least one of the circumferentially-spaced sides, and coupling the basesheet to the backbone assembly.

Abstract: A gas turbine engine exhaust nozzle includes a divergent section located aft of a convergent section and a throat therebetween. An exterior fairing is spaced radially outwardly of the divergent section. An ejector cooling air flowpath leads from an ejector cooling air inlet in an aft portion of the fairing to a cooling air ejector in the nozzle. An annular nozzle plenum may be disposed between the divergent section of the nozzle and the external fairing and be part of the ejector cooling air flowpath between the ejector cooling air inlet and the ejector. A plurality of divergent flaps and divergent seals in the divergent section may employ cooling air passages, such as slots, to serve as the ejector. The fairing may include a plurality of circumferentially adjacent exterior flaps and exterior seals and employ truncated ends of or apertures in the exterior seals as the ejector cooling air inlet.