Abstract: A gas turbine engine including a core having a compressor section fluidly connected to a combustor via a primary flowpath and a turbine section connected to the combustor via the core flow path. An assembly is disposed within the gas turbine engine and includes a first part connected to a second part via a radial stack joint. The first part includes a radially inward facing surface contacting a corresponding radially outward facing surface of the second part. A fastener protrudes through the first part and the second part and is configured to maintain the relative positions of the first part and the second part. A channel is disposed on at least one of the radially inward facing surfaces and is positioned between the fastener and a circumferential edge of the first part. The channel is connected to at least one cooling air source.

Abstract: A component for a gas turbine engine includes an airfoil, including therein a first cavity and a second cavity, and an outer spindle extending from the airfoil, the outer spindle including a first channel in fluid communication with the first cavity and a second channel in fluid communication with the second cavity.

Abstract: A gas turbine engine including a core having a compressor section fluidly connected to a combustor via a primary flowpath and a turbine section connected to the combustor via the core flow path. An assembly is disposed within the gas turbine engine and includes a first part connected to a second part via a radial stack joint. The first part includes a radially inward facing surface contacting a corresponding radially outward facing surface of the second part. A fastener protrudes through the first part and the second part and is configured to maintain the relative positions of the first part and the second part. A channel is disposed on at least one of the radially inward facing surfaces and is positioned between the fastener and a circumferential edge of the first part. The channel is connected to at least one cooling air source.

Abstract: A vane assembly may comprise a plurality of airfoils each extending between an inner platform and an outer platform, the plurality of airfoils comprising a nominal airfoil, a thickened airfoil, and a first thinned airfoil circumferentially adjacent to the thickened airfoil, wherein, the nominal airfoil has a first chord thickness, the thickened airfoil has a second chord thickness, and the thinned airfoil has a third chord thickness, wherein the second chord thickness is greater than the first chord thickness and the third chord thickness is less than the first chord thickness, and a member disposed within the thickened airfoil.

Abstract: A gas turbine engine component assembly including: a first component having a first surface and a second surface opposite the first surface; and a second component having a first surface, a second surface opposite the first surface of the second component, and an impingement slot extending from the second surface of the second component to the first surface of the second component, the second surface of the first component and the first surface of the second component defining a cooling channel therebetween in fluid communication with the impingement slot, wherein the impingement slot is in fluid communication with the second surface of the first component, wherein the impingement slot includes a slot tab configured to direct airflow into the cooling channel at least partially in a lateral direction parallel to the second surface of the first component such that a cross flow is generated in the cooling channel.

Abstract: Vanes for gas turbine engines are described. The vanes include an airfoil body extending between a first platform and a second platform, the airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side and at least one skin core cavity comprising a first skin core cavity, wherein the first skin core cavity defines a throughflow passage through the vane between the first platform and the second platform, wherein the at least one skin core cavity is defined between an external hot wall of the vane and an internal cold wall of the vane.

Abstract: A heat shield panel arrangement for use in a combustor of a gas turbine engine is disclosed. In various embodiments, the heat shield panel arrangement includes a first panel having a first rail, the first rail having a first distal tip and a first angled outer wall, the first distal tip having a first distal tip length; and a second panel having a second rail, the second rail having a second distal tip and a second angled outer wall, the second distal tip having a second distal tip length. The first distal tip length is less than the second distal tip length and the first angled outer wall is substantially parallel to the second angled outer wall with a gap formed therebetween.

Abstract: A turbine vane includes a platform, an airfoil, and a cooling arrangement. The platform has a first surface and a second surface. The first surface and the second surface each axially between a first platform end and a second platform end and each circumferentially extend between a first platform side and a second platform side. The airfoil is connected to the platform and defines a passage outlet that extends through the second surface. The cooling arrangement includes a third wall that is connected to the second surface of the platform. The third wall defines a first outlet and a second outlet.

Abstract: A vane assembly may comprise a plurality of airfoils each extending between an inner platform and an outer platform, the plurality of airfoils comprising a nominal airfoil, a thickened airfoil, and a first thinned airfoil circumferentially adjacent to the thickened airfoil, wherein, the nominal airfoil has a first chord thickness, the thickened airfoil has a second chord thickness, and the thinned airfoil has a third chord thickness, wherein the second chord thickness is greater than the first chord thickness and the third chord thickness is less than the first chord thickness, and a member disposed within the thickened airfoil.

Abstract: Baffles and airfoils for installation within gas turbine engines are provided. The baffles include a baffle side wall extending between a first end and a second end and a baffle divider within the baffle side wall, wherein the baffle divider and the baffle side wall define a first cavity and a second cavity. The first cavity is located at a first side at the first end and the second cavity is located at a second side at the first end and the baffle divider includes a spiral portion such that the first cavity is transitioned toward the second side and the second cavity is transitioned toward the first side.

Abstract: An exemplary gas turbine engine includes a turbine section positioned about an engine central longitudinal axis. The turbine section includes a component with a platform providing a lip, a rail extending radially from the platform and at an axial location spaced from an outer axial extension of the lip. An inner face of the rail and a surface of the platform at least partially provide a cavity. At least one opening extends from the inner face to an outer face of the rail opposite the inner face to provide fluid communication from the cavity to the lip.

Abstract: A gas turbine engine includes a first vane of a turbine section, a second vane of the turbine section, an impingement plate that includes a first impingement hole and a second impingement hole, and a pipe that supplies a fluid. The first vane includes a first platform. The second vane includes a second platform and a baffle. A first portion of the fluid traverses the first impingement hole and impinges on a first surface of the first platform. A second portion of the fluid traverses the second impingement hole and impinges on a second surface of the second platform. A third portion of the fluid bypasses the impingement plate and enters the baffle.

Abstract: Partitioned baffles and components for installation within gas turbine engines are provided. The partitioned baffles include a baffle body defining a first cavity, the baffle body having a first inlet and at least one first impingement hole arranged for flow to pass through the first cavity from the first inlet to the at least one first impingement hole and a partitioned channel installed within the baffle body and defining a second cavity, wherein the partitioned channel has a second inlet and at least one second impingement hole, wherein a flow through the second cavity flows from the second inlet to the at least one second impingement hole. The baffle body is configured to receive cooling air from a first source and the partitioned channel is configured to receive cooling air from a second source, wherein the first source is different from the second source.

Abstract: Components for gas turbine engines are provided. The components include a platform, a first airfoil extending from the platform, a first cavity located within the first airfoil, the first cavity includes a forward portion and an aft portion separated by a divider, a cover plate attached to the platform on a side opposite the airfoil, wherein a platform cavity is defined between the cover plate and the platform and wherein the first cavity is fluidly connected to the platform cavity through a first cavity inlet, and a first cavity separating rail dividing the platform cavity into a first platform forward cavity and a first platform aft cavity, wherein the first platform forward cavity is fluidly connected to the aft portion of the first cavity and the first platform aft cavity is fluidly connected to the forward portion of the first cavity.

Abstract: A variable vane system for a gas turbine engine includes an airfoil trunnion that extends through a vane platform such that the airfoil is rotatable with respect to the vane platform about an axis, the airfoil trunnion having an inlet configuration into the airfoil and a fixed airflow inlet mask to abut the airfoil trunnion, the fixed airflow inlet mask having a mask inlet configuration that modulates an airflow through the inlet configuration.

Abstract: Vanes for gas turbine engines are described. The vanes include an airfoil body extending between a first platform and a second platform, the airfoil body having a leading edge, a trailing edge, a pressure side, and a suction side and at least one skin core cavity comprising a first skin core cavity, wherein the first skin core cavity defines a throughflow passage through the vane between the first platform and the second platform, wherein the at least one skin core cavity is defined between an external hot wall of the vane and an internal cold wall of the vane.

Abstract: A heat shield panel arrangement for use in a combustor of a gas turbine engine is disclosed. In various embodiments, the heat shield panel arrangement includes a first panel having a first rail, the first rail having a first distal tip and a first angled outer wall, the first distal tip having a first distal tip length; and a second panel having a second rail, the second rail having a second distal tip and a second angled outer wall, the second distal tip having a second distal tip length. The first distal tip length is less than the second distal tip length and the first angled outer wall is substantially parallel to the second angled outer wall with a gap formed therebetween.

Abstract: A component for a gas turbine engine is disclosed. In various embodiments, the component includes an airfoil, including therein a first cavity and a second cavity, and an outer spindle extending from the airfoil, the outer spindle including a first channel in fluid communication with the first cavity and a second channel in fluid communication with the second cavity.

Abstract: An exemplary gas turbine engine includes a turbine section positioned about an engine central longitudinal axis. The turbine section includes a component with a platform providing a lip, a rail extending radially from the platform and at an axial location spaced from an outer axial extension of the lip. An inner face of the rail and a surface of the platform at least partially provide a cavity. At least one opening extends from the inner face to an outer face of the rail opposite the inner face to provide fluid communication from the cavity to the lip.

Abstract: A heat exchanger for a gas turbine engine including: a heat exchanger core having an inlet, an outlet, and a fluid passageway; an inlet pipe configured to provide hot airflow to the inlet; and an inlet flow conditioner fluidly connecting the outlet of the inlet pipe to the inlet of the heat exchanger core, the inlet flow conditioner including: a central attachment body; an outer shell located radially outward from the central attachment body, the outer shell including an inlet connected to the outlet of the inlet pipe, and an outlet opposite the inlet and connected to the inlet of the heat exchanger core; and a plurality of guide vanes extending from the central attachment body to the outer shell, the plurality of guide vanes being configured to impart a swirl upon airflow entering the inlet of the inlet flow conditioner and exiting the outlet of the inlet flow conditioner.