Abstract: A power turbine rotor locking assembly including a rotor shaft. The assembly also includes an aft rotor stage surrounding the rotor shaft. The assembly further includes a locknut torqued to the rotor shaft. The assembly yet further includes a plurality of torque resistance features located on an aft side of the aft rotor stage. The assembly also includes a tooling fixture having a plurality of corresponding torque resistance features, the corresponding torque resistance features selectively engageable with the torque resistance features of the aft rotor stage to resist circumferential movement during assembly and disassembly of the locknut with the rotor shaft.

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 active clearance control manifold assembly includes multiple arcuate manifold segments each having multiple circumferential channels axially spaced apart from one another. The circumferential channels include cooling holes facing radially inward. A tube at least partially circumscribes and fluidly interconnects the manifold segments.

Abstract: A liner panel for use in a combustor of a gas turbine engine, including a multiple of heat transfer augmentors located in at least one discrete area of the liner panel. A wall assembly within a gas turbine engine including a support shell, a liner panel mounted to the support shell via a multiple of studs and a multiple of heat transfer augmentors located in at least one discrete area on a cold side of the liner panel, each of the multiple of heat transfer augmentors define a height that extends about half the distance between a cold surface of the liner panel and the support shell.

Abstract: A liner panel for use in a combustor of a gas turbine engine, the liner panel includes a radiused edge blended into a hot side of the linear panel. A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a radiused edge blended into a hot side of the linear panel, the hot side including a thermal barrier coating. A method of manufacturing including casting a radiused edge blended into a hot side of a linear panel; and applying a thermal barrier coating to the hot side of the liner panel.

Abstract: A strut for a gas turbine engine includes an airfoil section extending in a spanwise direction between a first platform and a second platform, extending in a chordwise direction between a leading edge and trailing edge to define a chord length, and extending in a thickness direction between a first side and a second side to define a chord width. Exterior surfaces of the airfoil section define a leading portion between the leading edge and a widest location of the airfoil section relative to the thickness direction, and a trailing portion between the widest location and the trailing edge. The exterior surfaces establish a respective exterior contour for each span position between a 0% span position and a 100% span position. The exterior surfaces define a plurality of dimples in the leading portion.

Abstract: A combustor for use in a gas turbine engine has a combustor outer shell. A panel has an inner face which will face hot products of combustion, and a boss surrounding a feature, with the boss extending to an outer end. A spacing surface is spaced from the boss, and is at an outer position that is inward of the outer end of the boss. The spacing surface spaces the panel from the outer shell. A trough is intermediate the boss and the spacing surface. The trough extends to an outer end which is inward of the outer position of the spacing surface. A gas turbine engine is also disclosed.

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: A heat shield for a gas turbine engine includes a radial heat shield flange that extends in a circumferential direction and forms a ring. A plurality of bosses extend from a first axial side of the radial heat shield flange. There is a plurality of guide pins. One of the plurality of guide pins extends from a corresponding one of the plurality of bosses.

Abstract: A gas turbine engine sealing system includes a case including forward and aft hooks. The aft hook arranged radially outward of the forward hook. A blade outer air seal has a J-hook that receives the aft hook and includes an end received by the forward hook.

Abstract: An example case for a gas turbine engine includes a bearing compartment, a radially outer ring, and a radially inner ring. The radially outer ring and radially inner ring are connected by a plurality of circumferentially spaced apart struts and define an annular flow path therebetween. The radially inner ring includes a first flange that extends radially outward from the bearing compartment. The first flange includes an undulation that extends axially away from the plurality of struts, and defines a load path between the bearing compartment and the radially outer ring.

Abstract: A liner panel for use in a combustor of a gas turbine engine, the liner panel includes a radiused gate blended into a hot side of the liner panel. A combustor for a gas turbine engine including a liner panel mounted to a support shell via a multiple of studs, the liner panel including a radiused gate blended into a hot side of the liner panel, the hot side including a thermal barrier coating. A method of manufacturing including casting a radiused gate tangentially cast into a hot side of a liner panel; and applying a thermal barrier coating to the hot side of the liner panel over the remnant of the radiused gate.

Abstract: A combustor seal for use in a combustor of a gas turbine engine including a seal with a multiple of slots that correspond with a multiple of 1st HPT vanes. A method of cooling within a gas turbine engine including communicating cooling air through combustor seal toward each of a multiple of 1st HPT vanes.

Abstract: A power turbine rotor locking assembly including a rotor shaft. The assembly also includes an aft rotor stage surrounding the rotor shaft. The assembly further includes a locknut torqued to the rotor shaft. The assembly yet further includes a plurality of torque resistance features located on an aft side of the aft rotor stage. The assembly also includes a tooling fixture having a plurality of corresponding torque resistance features, the corresponding torque resistance features selectively engageable with the torque resistance features of the aft rotor stage to resist circumferential movement during assembly and disassembly of the locknut with the rotor shaft.

Abstract: A turbine blade assembly for a gas turbine engine is provided. The turbine blade assembly having: a rotor; a plurality of turbine blades mounted to the rotor, each of the plurality of turbine blades having a forward tab that contacts a front face of the rotor and an aft tab; a plurality of tabs that extend from an aft surface of the rotor; and a retaining ring that engages each aft tab of the plurality of turbine blades and each of the plurality of tabs that extend from the aft surface of the rotor.

Abstract: A turbine section for a gas turbine engine includes blade outer air seals and stator vanes that provide a core flow path. A turbine case supports blade outer air seals and stator vanes. An annular cavity is provided between an interior surface of the turbine case and the blade outer air seals and the stator vanes. A hole extends through the turbine case from an exterior surface to the interior surface. The annular cavity extends axially to an exit. A manifold circumscribes the exterior surface of the turbine case and provides an annular space therebetween. The annular space is in fluid communication with the exit of the annular cavity via the hole.

Abstract: A gas turbine engine includes a fan, a compressor section, a combustor, a fan drive gear system, and a turbine section coupled to drive the fan through the gear system. The combustor includes an annular outer shell and an annular inner shell that define an annular combustion chamber. There is a bulkhead in the annular combustion chamber, and an annular heat shield is mounted on the bulkhead. The annular heat shield includes a segment that has a forward face and an aft face, a circumferential outer side and a circumferential inner side, a central orifice between the forward face and the aft face, a lip projecting from the forward face around the central orifice, a rail projecting from the forward face, and a plurality of through-holes between the rail and the lip. The rail contacts the bulkhead to define a cavity bounded by the rail, the lip, and the bulkhead.

Abstract: A flange heat shield is provided. The flange heat shield may comprise an annular body having a forward end opposite an aft end. The forward end of the flange heat shield may comprise a radial snap configured to interface with an inner surface of an engine case. The inner surface of the engine case may comprise an airflow ramp forward of the radial snap. The aft end of the flange heat shield may comprise a plurality of notches defining voids on the aft end. The aft end of the flange heat shield may also comprise mounting tabs configured to couple the flange heat shield to the engine case.

Abstract: A heat shield for a gas turbine engine includes a radial heat shield flange that extends in a circumferential direction and forms a ring. A plurality of bosses extend from a first axial side of the radial heat shield flange. There is a plurality of guide pins. One of the plurality of guide pins extends from a corresponding one of the plurality of bosses.

Abstract: An example case for a gas turbine engine includes a bearing compartment, a radially outer ring, and a radially inner ring. The radially outer ring and radially inner ring are connected by a plurality of circumferentially spaced apart struts and define an annular flow path therebetween. The radially inner ring includes a first flange that extends radially outward from the bearing compartment. The first flange includes an undulation that extends axially away from the plurality of struts, and defines a load path between the bearing compartment and the radially outer ring.