Abstract: A gas-turbine engine is provided. The gas-turbine engine comprises a high pressure turbine with an aft blade platform. A static structure may be disposed aft of the high pressure turbine and proximate a cavity defined by the aft blade platform. A vane of the static structure may have a vane platform with a shaped tip extending into the cavity.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A first vane stage assembly and a method are also disclosed.

Abstract: A heat shield panel for a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side defining a first surface having an outer perimeter, a cold side defining a second surface spaced from the first surface, a rail member disposed on the cold side proximate a first portion of the outer perimeter, the rail member having an outer wall and an inner wall, an undercut portion within the second surface proximate the inner wall of the rail member and an orifice extending through the rail member, from the inner wall to the outer wall.

Abstract: An investment casting method includes providing a stock investment casting core, bending the stock investment casting core to thereby form a production investment casting core that conforms to a design cooling passage shape, and casting an alloy around the production investment casting core to form a cast article.

Abstract: A method of assembling a mid-turbine frame includes engaging at least two tabs in a radially inner end of at least one spoke in a groove in an inner frame case. The at least one spoke is secured to the inner frame case with a central bolt extending through the inner frame case into the at least one spoke.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A plurality of vane cooling passages extending through a forward flange of the first vane stage to the radial inner surface for communicating cooling airflow into the core flow path forward of the leading edge. A seal assembly is disposed between the combustor and the forward flange of the vane stage. The seal includes a plurality of circumferentially spaced apart slots and a plurality of seal cooling passages that extend from the radially outer surface to the radially inner surface at each of the plurality of circumferentially spaced apart slots.

Abstract: A method of manufacturing a heat shield panel including pouring melted wax into a negative pattern of the heat shield panel, the heat shield panel including first pin fins with rounded tops and second pin fins with flat tops; allowing the wax to solidify to form a positive pattern of the heat shield panel; removing the positive pattern from the negative pattern by using an ejector rod to push the positive pattern away from the negative pattern at the flat top of each of the one or more second pin fins; coating the positive pattern with a ceramic; melting the positive pattern away from the ceramic, the ceramic having a cavity forming a second negative pattern of the heat shield panel; pouring melted metal into the cavity; allowing metal in the cavity to cool to form the heat shield panel; and removing the ceramic from the heat shield panel.

Abstract: A heat shield panel for use in a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side defining a first surface having an outer perimeter, a cold side defining a second surface spaced from the first surface, a rail member disposed on the cold side proximate a first portion of the outer perimeter, the rail member having an outer wall and an inner wall, and an orifice extending through the rail member, from the inner wall to the outer wall, the orifice having a cross sectional area that increases from the inner wall to the outer wall.

Abstract: A method of manufacturing a heat shield panel including pouring melted wax into a negative pattern of the heat shield panel, the heat shield panel including first pin fins with rounded tops and second pin fins with flat tops; allowing the wax to solidify to form a positive pattern of the heat shield panel; removing the positive pattern from the negative pattern by using an ejector rod to push the positive pattern away from the negative pattern at the flat top of each of the one or more second pin fins; coating the positive pattern with a ceramic; melting the positive pattern away from the ceramic, the ceramic having a cavity forming a second negative pattern of the heat shield panel; pouring melted metal into the cavity; allowing metal in the cavity to cool to form the heat shield panel; and removing the ceramic from the heat shield panel.

Abstract: A gas turbine engine component assembly including: a first component having a first surface and a second surface opposite the first surface; a second component having a first surface and a second surface opposite the first surface of the second component; and a cooling circuit formed in the second component, the cooling circuit including an inlet at the second surface of the second component, an outlet at the first surface of the second component, an inward surface, and an outward surface, the inward surface and the outward surface being in a facing spaced relationship extending from the inlet to the outlet defining an airflow passageway therebetween, wherein the airflow passageway includes a parallel portion that is oriented about parallel to at least one of the first surface of the second component and the second surface of the second component.

Abstract: A heat shield panel for a gas turbine engine combustor may comprise a perimetrical rail extending at least partially around a perimeter of the heat shield panel and defining a main backside cavity, a plurality of air admittance holes, each air admittance hole extending along a hole axis, and a midrail dividing the main backside cavity into a first backside cavity and a second backside cavity, wherein at least a portion of the midrail is axially offset from at least one of the hole axes.

Abstract: A combustor for use in a gas turbine engine, the combustor enclosing a combustion chamber having a combustion area. The combustor includes: a shell having a kink; and a kinked heat shield panel in facing spaced relationship with the shell, the kinked heat shield panel including a kink located proximate the kink in the shell, wherein the kinked heat shield panel further includes a first surface, a second surface opposite the first surface, and a mounting stud located proximate the kink of the kinked heat shield panel and extending away from the second surface.

Abstract: An aerodynamic component of a gas turbine engine is provided and is fittable to a shell having a shell shape. The aerodynamic component includes a body having a component shape initially deviating from the shell shape prior to an assembly operation in which the aerodynamic component is to be fit to the shell. Deviation of the component shape from the shell shape aids in an establishment of a final desired shape of the aerodynamic component following the assembly operation.

Abstract: An investment casting method includes providing a stock investment casting core, bending the stock investment casting core to thereby form a production investment casting core that conforms to a design cooling passage shape, and casting an alloy around the production investment casting core to form a cast article.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a first plurality of openings and the first vane stage includes a second plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A first vane stage assembly and a method are also disclosed.

Abstract: A seal system is provided. The seal system may comprise a first duct having an annular geometry, a second duct overlapping the first duct in a radial direction, and a seal disposed between the first duct and the second duct. The seal may comprise a groove defined by the first duct and a piston configured to slideably engage the groove.

Abstract: A heat shield panel for a combustor of a gas turbine engine including a panel body having a first surface and a second surface. The second surface being configured to be oriented toward a combustor liner of the combustor. The heat shield further includes a plurality of first pin fins projecting from the second surface of the panel body. Each of the plurality of first pin fins has a rounded top opposite the second surface. The heat shield further includes a plurality of second pin fins projecting from the second surface of the panel body. Each of the plurality of second pin fins has a flat top opposite the second surface. The plurality of second pin fins are intermittently spaced amongst the plurality of first pin fins. The plurality of second pin fins are organized in a uniform distribution across the second surface of the heat shield panel.

Abstract: Aspects of the disclosure are directed to a seal comprising: a first fitting configured to couple to a first disk, and a curvic joint including curvic teeth, where a first distance between the first fitting and the curvic teeth is equal to or greater than a first thickness of the first disk. In some embodiments, an engine comprises: a compressor disk, a turbine disk, and a seal including: a first fitting configured to couple to the turbine disk, a second fitting configured to couple to the compressor disk, and a curvic joint including curvic teeth, where a first axial distance between the first fitting and the curvic teeth is greater than or equal to a first radial thickness of the turbine disk.

Abstract: A gas turbine engine includes a combustor. A turbine section is in fluid communication with the combustor. The turbine section includes a first vane stage aft of the combustor. A seal assembly is disposed between the combustor and the first vane stage. The seal assembly includes a plurality of openings communicating cooling airflow into a gap between an aft end of the combustor and the first vane stage. A combustor assembly and method are also disclosed.

Abstract: A combustor panel may comprise a panel wall comprising a proximal surface and a distal surface. A standoff may be formed over the distal surface of the panel wall. An aperture may be formed through the standoff and may extend from a face of the standoff to the proximal surface of the panel wall. The face of the first standoff may be oriented at an angle between 90 degrees and 160 degrees relative to a plane parallel to the distal surface of the panel wall.