Abstract: A ducted heat exchanger system for a gas turbine engine includes an additive manufactured heat exchanger core with a contoured external and/or internal geometry. A method of additively manufacturing a heat exchanger for a gas turbine engine includes additively manufacturing a core of a heat exchanger to set a ratio of local surface area to flow area to control a pressure drop per unit length along the core.

Abstract: A retaining assembly may comprise a bracket assembly having a first end and a second end. The first end may be configured to mount to a clamping band. The retaining assembly may further comprise a retention component. The retention component may have a clevis coupled to the second end. The retaining assembly may be coupled to an otherwise loose external component on a gas turbine engine. The retaining assembly may keep the otherwise loose external component fixed to the gas turbine engine during maintenance operations and prevent the otherwise loose external component from being lost or misplaced.

Abstract: A slider seal assembly for a gas turbine engine, comprising a housing comprising an outer surface and an inner surface, and a recessed opening between the inner surface and the outer surface, wherein the inner surface of the housing includes a flat portion configured to match a receiver on an inner wall of a casing of the gas turbine engine; a seal plate received within the recessed opening and moveable relative to the housing; and a retaining ring configured to retain the seal plate adjacent the housing, the retaining ring includes a plurality of tabs configured to retain the seal plate adjacent the housing.

Abstract: A disclosed turbofan engine includes a gas generator section for generating a gas stream flow. A speed reduction device is driven by the power turbine. A propulsor section includes a fan driven by the power turbine through the speed reduction device at a second speed lower than the first speed for generating propulsive thrust as a mass flow rate of air through a bypass flow path. The fan includes a tip diameter greater than about fifty (50) inches and an Engine Unit Thrust Parameter (“EUTP”) defined as net engine thrust divided by a product of the mass flow rate of air through the bypass flow path, a tip diameter of the fan and the first rotational speed of the power turbine is less than about 0.30 at a take-off condition.

Abstract: A method of making a light weight component is provided. The method including the steps of: forming a metallic foam core into a desired configuration; and applying an external metallic shell to an exterior surface of the metallic foam core after it has been formed into the desired configuration.

Abstract: A fan assembly for a gas turbine engine includes a fan including a plurality of fan blades. Each fan blade extends radially outwardly from a fan hub to a blade tip. The plurality of blade tips define a fan diameter. A nacelle surrounds the fan and defines a fan inlet upstream of the fan, relative to an airflow direction into the fan. The nacelle has a forwardmost edge defining an inlet length from the forwardmost edge to a leading edge of a fan blade of the plurality of fan blades. A ratio of inlet length to fan diameter is between 0.20 and 0.45. A nacelle inner surface defines a nacelle flowpath. The nacelle flowpath has a convex throat portion and a concave diffusion portion between the throat portion and the leading edge of the fan blade at a bottommost portion of the nacelle.

Abstract: A stator vane may comprise an airfoil extending between a first platform and a second platform, the airfoil including a core extending relatively orthogonal to the first platform and the second platform, at least one of the first platform or second platform comprising a frustic load transmission feature, wherein the frustic load transmission feature comprises at least a first angular surface disposed proximate a platform edge, wherein the first angular surface is defined by a non-orthogonal angle ? with respect to an outer platform surface.

Abstract: An airfoil may comprise a root and an airfoil body radially outward of the root. The airfoil body may define a first cooling chamber and a second cooling chamber. A first passage may be defined within the root and configured to direct a first airflow radially outward through the root into the first cooling chamber. A second passage may be defined within the root and configured to direct a second airflow radially outward through the root and into the second cooling chamber.

Abstract: A seal assembly for a turbine engine extends along a center axis and includes a seal support and a seal carrier configured to translate relative the seal support. A seal is connected to the seal carrier and a spring is disposed between the seal support and the seal carrier. A spring carrier is disposed between the spring and the seal support. A first end of the spring is connected to the spring carrier and the spring carrier is connected to the seal support. The spring includes a second end disposed opposite the first end of the spring. The second end of the spring contacts the annular seal carrier.

Abstract: A method for heat treating a superalloy component includes heating a superalloy component to a first temperature, cooling the superalloy from the first temperature to a second temperature at a first cooling rate in a furnace, and cooling the superalloy component from the second temperature to a final temperature at a second cooling rate. The second cooling rate is higher than the first cooling rate.

Abstract: A gas turbine engine comprises a housing having an inlet leading to a fan rotor. A bypass door is mounted upstream of the inlet to the fan rotor, and is moveable away from a non-bypass position to a bypass position to selectively bypass boundary layer air vertically beneath the engine. An aircraft is also disclosed.

Abstract: A system for direct writing a conformal or free-form magnet includes producing a magnetic ink compound by mixing an photopolymer base and a magnetic material powder, extruding a layer of the magnetic ink compound through a nozzle on to a target substrate to form a layer, soft-curing the layer by exposing the layer to a first light, repeatedly extruding a layer and soft-curing the layer, then post-curing the magnet by exposing the magnet to a second light and an elevated temperature.

Abstract: A blade outer air seal assembly may comprise a blade outer air seal segment and a blade outer air seal support coupled to the blade outer air seal segment. The blade outer air seal segment may comprise a forward rail and an aft rail. The forward rail may be castellated. The blade outer air seal support may comprise an aft flange and a first forward flange. A pin may be disposed through the aft rail, the aft flange, the first forward flange, and the forward rail.

Abstract: A blade outer air seal retention assembly may comprise a forward retention ring and an aft retention ring coupled to the forward retention ring. The forward retention ring may comprise a first shiplap flange extending aft from the forward retention ring. The aft retention ring may comprise a second shiplap flange extending forward from the aft retention ring.

Abstract: A shroud assembly may comprise a first ring, a second ring aft of the first ring, and a shroud disposed between the first ring and the second ring. The shroud may comprise a plurality of circumferentially adjacent shroud segments. Each shroud segment of the plurality of circumferentially adjacent shroud segments may extend axially from the first ring to the second ring.

Abstract: A monolithic airflow testing mold suitable for mating with a workpiece includes a platform portion formed from a first material having a first hardness. The mold further includes a workpiece mating portion surrounding by the platform portion. The mating portion includes a support region abutting the platform portion and formed from a second material having a second hardness, and a sealing surface positioned above the support region and formed from a third material having a third hardness. The sealing surface is configured to mate with and extend into a hollow portion of the workpiece. The first hardness is different from at least one of the second and third hardnesses.

Abstract: A lubrication system for use with a gas turbine engine includes a first reservoir for containing a lubricant. The first reservoir includes a first discharge passage through which the lubricant is flowable in a first direction. A second reservoir contains the lubricant. The second reservoir includes a second discharge passage through which the lubricant is flowable in a second direction. The first direction is generally opposite to the second direction. A first pump pumps the lubricant from the first reservoir. A second pump pumps the lubricant from the second reservoir. A manifold distributes the lubricant to a component. The lubricant from the first pump and the second pump flows into the manifold and exits the manifold through a manifold discharge.

Abstract: A method for monitoring an additive manufacturing process during fabrication of a component part is disclosed. In various embodiments, the method includes the steps of selecting a sensing matrix; orienting a sensor toward a surface of the component part; generating a discrete time signal, based on data obtained from the sensor, the discrete time signal being representative of a process condition of the component part while the component part is undergoing the additive manufacturing process; compressing the discrete time signal using the sensing matrix to form a compressed measurement signal; and storing the compressed measurement signal in a storage device while the component part is undergoing the additive manufacturing process. In various embodiments, selecting the sensing matrix comprises selecting a basis function. In various embodiments, the basis function is determined using a random time sampling.

Abstract: An injector of a radial fuel injection system for a combustor of a gas turbine engine includes a swirler; and a fuel nozzle located within the swirler, the fuel nozzle located within the swirler, the fuel nozzle operable to provide a biased radial fuel distribution within the swirler. A method of controlling a fuel flow to a combustor of a gas turbine engine includes selectively controlling a biased radial fuel distribution within a swirler of a radial fuel injection system.

Abstract: An end effector for a vacuum system is disclosed. In various embodiments, the end effector includes a central duct having a first end configured for connection to a vacuum source and a second end defining a central duct inlet; a base member proximate the second end of the central duct; and a plurality of passages extending from an outer surface of the base member to an inner surface of the central duct, the plurality of passages characterized by a passage axis having an axial vector component.