Abstract: A combustor liner for a gas turbine engine, the combustor liner including a panel configured to at least partially define a combustion chamber. The combustor liner further includes a shell configured to mount to the panel and form a gap between the panel and the shell. The panel includes a stud and a plurality of a stand-off pins proximate to the stud defining a cavity therebetween. The shell includes a plurality of angled impingement holes located away from the cavity but extending through the shell at an orientation such that cooling air passing through the angled impingement holes is directed towards the cavity between adjacent stand-off pins and at an acute angle relative to the stud.

Abstract: A compressor case for a gas turbine engine includes an annular body that extends circumferentially around a center axis and extends axially along the center axis. A first bleed manifold is formed on an outer surface of the annular body and encloses a first plenum. A second bleed manifold is formed on the outer surface of the annular body and is axially aft of the first bleed manifold. The second bleed manifold encloses a second plenum. A bleed inlet extends through the annular body and into the first bleed manifold. Cooling passages are formed in the annular body, and each of the cooling passages extends from the first plenum to the second plenum and fluidically connects the first plenum to the second plenum.

Abstract: A gas turbine engine includes a first spool, a second spool, a primary combustor, and a diffuser. The first spool includes a first compressor rotationally driven by a first turbine via a first shaft. The second spool includes a second compressor driven by a second turbine via a second shaft. The first compressor, the diffuser, and the primary combustor are arranged in series to provide a compressed airflow discharged from the first compressor to the primary combustor via the diffuser, which includes walls that diverge towards the primary combustor. The second compressor is fluidly coupled to the diffuser to receive at least a portion of the compressed airflow from the diffuser. The second turbine is fluidly coupled to the diffuser to discharge an expanded airflow to the diffuser.

Abstract: A rotor blade for a gas turbine engine including an inner ply layer group includes a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter.

Abstract: A furnace for removing a molybdenum-alloy refractory metal core through sublimation comprising a retort furnace having an interior; a sublimation fixture insertable within the interior of the retort furnace, the sublimation fixture configured to receive at least one turbine blade having the molybdenum-alloy refractory metal core; a flow passage thermally coupled to the retort furnace configured to heat a fluid flowing through the flow passage and deliver the fluid to the molybdenum-alloy refractory metal core causing sublimation of the molybdenum-alloy refractory metal core.

Abstract: A nacelle assembly of a gas turbine engine includes an annular structure defining a central axis, and having a radially inward surface and a radially outward surface, the radially inward surface at least partially defining a bypass duct. An aft portion of the radially inward surface at least partially defines an axially extending convergent-divergent exit nozzle. A secondary nozzle flap is radially spaced from the aft portion of the radially inward surface. The secondary nozzle flap and the aft portion of the radially inward surface define a secondary bypass duct therebetween. The secondary nozzle flap is operably connected to the annular structure such that the secondary nozzle flap is selectably movable relative to the aft portion of the radially inward surface, thereby changing a cross-sectional area of a secondary bypass duct exit.

Abstract: A control system for limiting a power turbine torque of a gas turbine engine is disclosed. In various embodiments, the control system includes an engine control module configured to output an effector command signal to a gas generator of the gas turbine engine; a power turbine governor module configured to output to the engine control module a power turbine torque request signal; and a power turbine torque limiter module configured to output to the power turbine governor module a power turbine speed rate signal to limit a power turbine speed overshoot of the gas turbine engine.

Abstract: A gas turbine engine includes an engine spool with a turbine section, a propeller, a motor, and a variable drive arrangement. The variable drive arrangement mechanically couples the engine spool to the propeller, and includes a first gear connection rotationally coupled to the propeller, a second gear connection rotationally coupled to the engine spool, and a sprag clutch rotationally coupled to the first gear connection. The sprag clutch is configured to passively decouple the motor from the first gear connection when a torque input from the engine spool exceeds a torque input from the motor, and is configured to passively couple the motor to the first gear connection when the torque input from motor exceeds a torque input from the engine spool.

Abstract: A nickel-phosphorus alloy coating comprising a substrate having a surface; a nickel phosphorus alloy coating plated to the surface, the nickel phosphorus alloy consisting of phosphorus from 15.0 wt. percent to 20.9 wt. percent.

Abstract: A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis, the rotor disk comprising a slot within a rim of the disk and a spacer within the slot, the spacer having a clearance feature sized to permit rotation of a rotor blade from an insertion position to an installed position.

Abstract: A vane arm connection system for a gas turbine engine includes a vane stem having a head with flat contact surfaces; a vane arm having a claw, the claw having opposed arms having inwardly facing surfaces engaging the flat contact surfaces of the head; and a torque transfer member having a body defining an opening for engaging the flat contact surfaces of the head of the vane stem, and at least one arm extending from the body to contact the claw, whereby load from torque is transferred away from the inwardly facing surfaces.

Abstract: A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis; and a damper pivot support that extends from a rim of the rotor disk, the damper pivot support comprising an aperture that defines a pivot axis parallel to the axis. A method to dampen a rotor blade of a gas turbine engine, includes independently pivoting a first damper and a second damper of a damper assembly about the pivot axis to respectively contact a first blade and a second blade in response to centrifugal force below a platform of the respective first blade and second blade.

Abstract: A rotor blade for a gas turbine engine including an inner ply layer group that at least partially defines a base of a root region; a platform around the inner ply layer group that at least partially defines the base; and a platform over-wrap around the platform, the platform over-wrap at least partially defines the base.

Abstract: A rotor stack for a gas turbine engine includes a first rotor disk with a first rotor spacer arm, the first rotor spacer arm having a first flange with an outboard flange surface and an inboard flange surface, a first hole along an axis through the first flange, the first hole having a counterbore in the outboard flange surface; a second rotor disk with a web having a second hole along the axis; a third rotor disk with a third rotor spacer arm, the third rotor spacer arm having a third flange with an outboard flange surface and an inboard flange surface, a third hole along the axis through the third flange, the third hole having a counterbore in the inboard flange surface; and a bushing with a tubular body and a flange that extends therefrom, the tubular body comprising at least one axial groove along an outer diameter thereof, the bushing extends through the first hole, the second hole and partially into the counterbore in the inboard flange surface of the third hole.

Abstract: A method for producing a metal boride powder includes producing a bonding gas stream from a first powder in a first fluidizing bed reactor, delivering the bonding gas stream to a second fluidized bed reactor through a conduit fluidly connecting the first and second fluidized bed reactors, fluidizing a second powder in the second fluidized bed reactor, mixing the second powder with the bonding gas stream such that a metal boride or boron-doped powder is formed.

Abstract: A vane ring for a gas turbine engine includes a contoured stop that extends from the static flowpath wall, the contoured stop being of an airfoil shape such that a first side of the contoured stop matches a portion of a first side of the first variable vane along a chord length when the first variable vane is pivoted about the longitudinal axis to a first position.

Abstract: A diffusion barrier coating on a nickel-based alloy substrate comprising the diffusion barrier being coupled to the substrate between the substrate and a composite material opposite the substrate, wherein the diffusion barrier comprises a nickel cobalt and chromium-aluminum-yttria powder material.

Abstract: A diffusion barrier coating on a nickel-based alloy substrate comprising the diffusion barrier being coupled to the substrate between the substrate and a composite material opposite the substrate, wherein the diffusion barrier comprises a nickel phosphorus alloy material.

Abstract: A side plate seal assembly for a gas turbine engine includes a multiple of non-metallic side plate seals that are arranged about an axis of the gas turbine engine to form a full hoop seal, each of the multiple of side plate seals comprise a retention surface and a knife edge seal surface that extends at an angle therefrom.

Abstract: An inspection feature for an adhesive bond portion between parts of a component comprising the inspection feature extending from a surface of the part of the component a predetermined bond line thickness, the inspection feature having a density configured to imitate a defect of the adhesive bond portion responsive to a non-destructive inspection scan.