Abstract: A radial inlet body of a gas turbine engine compressor is provided. The radial inlet body includes plenums arranged around opposite sides of an inlet portion defining a compressor inlet. The plenums include outer surfaces with increasing radii of curvature and a flow diverter. The flow diverter is disposed aside the inlet portion opposite a duct. The plenums are configured to direct flows from the duct around the inlet portion. The flow diverter and the plenums are configured to cooperatively encourage uniform distribution of the flows. The flow diverter includes curvilinear sides that tangentially register with and smoothly lead to the outer surfaces of the plenums and a cusp where the curvilinear sides meet. The radial inlet body further includes a front sidewall supporting the inlet portion and an aft sidewall opposite the front sidewall. The front and aft sidewalls are angled toward one another along the plenums.

Abstract: A radial inlet body of a gas turbine engine compressor is provided. The radial inlet body includes plenums arranged around opposite sides of an inlet portion defining a compressor inlet. The plenums include outer surfaces with increasing radii of curvature and a flow diverter. The flow diverter is disposed aside the inlet portion opposite a duct. The plenums are configured to direct flows from the duct around the inlet portion. The flow diverter and the plenums are configured to cooperatively encourage uniform distribution of the flows. The flow diverter includes curvilinear sides that tangentially register with and smoothly lead to the outer surfaces of the plenums and a cusp where the curvilinear sides meet. The radial inlet body further includes a front sidewall supporting the inlet portion and an aft sidewall opposite the front sidewall. The front and aft sidewalls are angled toward one another along the plenums.

Abstract: A plenum assembly for a compressor has a plenum housing defining a radially inward entry to a compressor and an outer circumference, and at least two open areas in the outer circumferences which are defined by at least two plenum flanges; an inlet duct connected to a first plenum flange of the at least two plenum flanges; and a panel having a flow control feature, wherein the panel is connected to a second plenum flange of the at least two plenum flanges.

Abstract: A non-bypass gas turbine engine is provided that includes compressor, combustor, and turbine sections, an engine compartment enclosure, an exhaust nozzle, and a secondary exhaust nozzle. The engine compartment enclosure is disposed radially outside of the compressor, combustor, and turbine sections and defines a first annular region outside of the compressor, combustor, and turbine sections. The exhaust nozzle is disposed downstream of the turbine section and is configured to receive the core gas flow exiting the turbine section. The exhaust nozzle includes a wall panel with a plurality of lobes that extend between forward and aft ends of the exhaust nozzle. Each lobe has a lobe height that increases in a direction from the forward end to the aft end of the exhaust nozzle. The lobes de-swirl the core gas flow exiting the turbine section and draw air from the first annular region during operation of the engine.

Abstract: A system for an aircraft includes an aircraft engine having a central axis with a flowpath projecting into the aircraft engine from an airflow inlet. An inlet plenum at the airflow inlet extends between a front wall and a rear wall along the central axis. An inlet guard is arranged at the airflow inlet and extends across the flowpath. The inlet guard includes a first screen extending circumferentially about the central axis and axially from the front wall to the rear wall, and a second screen radially outward of the first screen and axially overlapping the first screen. An actuator is operable to move the first screen and the second screen relative to one another in one or more of a radial, axial and circumferential direction to shed ice from the inlet guard.

Abstract: A plenum assembly for a compressor has a plenum housing defining a radially inward entry to a compressor and an outer circumference, and at least two open areas in the outer circumferences which are defined by at least two plenum flanges; an inlet duct connected to a first plenum flange of the at least two plenum flanges; and a panel having a flow control feature, wherein the panel is connected to a second plenum flange of the at least two plenum flanges.

Abstract: An exhaust mixer assembly for a gas turbine engine includes a core passage extending along a central engine axis for directing a core gas flow. An outer annular passage coaxially surrounds the core passage for directing a bypass gas flow. An exhaust mixer communicating with the core passage and the outer annular passage has an upstream end, a downstream end and an annular wall extending therebetween. A plurality of protrusions extend axially from a downstream end of the annular wall to form a jagged trailing edge of the exhaust mixer. Each of the plurality of protrusions extend separately from the downstream end of the annular wall from the other of the plurality of protrusions. At least one of the plurality of protrusions is twisted about a protrusion axis extending through the at least one of the plurality of protrusions, the protrusion axis parallel to the central axis.

Abstract: An exhaust duct of an aircraft engine includes an annular inlet conduit having an inlet central axis, and at least two outlet conduits in flow communication with the inlet conduit. The at least two outlet conduits are located non-parallel to the inlet central axis. Each of the at least two outlet conduits include an outlet port defining a distal end of each of the two outlet conduits. At least one of the outlet ports is non-circular in cross-sectional shape.

Abstract: A system is provided for an aircraft. This aircraft system includes a gas turbine engine, a flowpath and an inlet guard. The gas turbine engine includes a compressor section. The flowpath projects longitudinally into the gas turbine engine from an airflow inlet and longitudinally through the compressor section. The inlet guard extends across the flowpath longitudinally upstream of the compressor section. The inlet guard extends circumferentially about an axis. The inlet guard includes a first screen and a second screen. The first screen includes a plurality of first perforations with a first perforation size. The second screen is circumferentially adjacent the first screen about the axis. The second screen includes a plurality of second perforations with a second perforation size that is different than the first perforation size.

Abstract: A system is provided for an aircraft. This aircraft system includes an aircraft engine, an inlet guard and an actuation system. The aircraft engine has a flowpath projecting into the aircraft engine from an airflow inlet. The inlet guard is arranged at the airflow inlet and extends across the flowpath. The inlet guard includes a first screen, a second screen and a plurality of perforations. The first screen is adjacent and overlaps the second screen. The perforations project through the inlet guard and are formed by the first screen and the second screen. The actuation system is configured to move the first screen along the second screen. The movement of the first screen along the second screen changes a geometry of each of the perforations.

Abstract: A system for an aircraft includes an aircraft engine having a central axis with a flowpath projecting into the aircraft engine from an airflow inlet. An inlet plenum at the airflow inlet extends between a front wall and a rear wall along the central axis. An inlet guard is arranged at the airflow inlet and extends across the flowpath. The inlet guard includes a first screen extending circumferentially about the central axis and axially from the front wall to the rear wall, and a second screen radially outward of the first screen and axially overlapping the first screen. An actuator is operable to move the first screen and the second screen relative to one another in one or more of a radial, axial and circumferential direction to shed ice from the inlet guard.

Abstract: A system for an aircraft includes an aircraft engine having a central axis with a flowpath projecting into the aircraft engine from an airflow inlet. An inlet plenum at the airflow inlet extends between a front wall and a rear wall along the central axis. An inlet guard is arranged at the airflow inlet and extends across the flowpath. The inlet guard includes a first screen extending circumferentially about the central axis. The first screen has a first screen axial width L1 extending from the front wall to the rear wall. The inlet guard further includes a second screen extending at least partially circumferentially about the central axis. The second screen is disposed radially outward of the first screen and axially overlaps the first screen. The second screen has a second screen axial width L2 less than the first screen axial width L1.

Abstract: A method of operation is provided during which rotation of a propulsor rotor of an aircraft is driven using mechanical power output from a powerplant. The powerplant includes a heat engine and an electric machine. The heat engine provides a first portion of the mechanical power. The electric machine provides a second portion of the mechanical power. An operational temperature of the heat engine is regulated by controlling the first portion of the mechanical power generated by the heat engine and the second portion of the mechanical power generated by the electric machine.

Abstract: An anti-IL-23 antibody, including isolated nucleic acids that encode at least one anti-IL-23 antibody, vectors, host cells, transgenic animals or plants, and methods of making and using thereof have applications in diagnostic and/or therapeutic compositions, methods and devices.

Abstract: An exhaust mixer assembly for a gas turbine engine includes a core passage extending along a central engine axis for directing a core gas flow. An outer annular passage coaxially surrounds the core passage for directing a bypass gas flow. An exhaust mixer communicating with the core passage and outer annular passages has an upstream end, a downstream end and an annular wall extending therebetween. The annular wall defines a mixing plane at a trailing edge thereof at the downstream end at which the core and bypass gas flows are mixed. The annular wall has a radially inner surface and a radially outer surface. A plurality of protrusions are disposed on and extend laterally along and radially from at the radially inner surface and/or the radially outer surface. The plurality of protrusions are arranged one alongside another on the radially inner surface and/or the radially outer surface.

Abstract: A human anti-IL-23p19 antibody, including isolated nucleic acids that encode at least one anti-IL-23p19 antibody, vectors, host cells, and methods of making and using thereof have applications in diagnostic and/or therapeutic compositions, methods and devices.

Abstract: An exhaust duct of an aircraft engine includes an annular inlet conduit having an inlet central axis, and at least two outlet conduits in flow communication with the inlet conduit. The at least two outlet conduits are located non-parallel to the inlet central axis. Each of the at least two outlet conduits include an outlet port defining a distal end of each of the two outlet conduits. At least one of the outlet ports is non-circular in cross-sectional shape.

Abstract: An exhaust mixer assembly for a gas turbine engine includes a core passage extending along a central engine axis for directing a core gas flow. An outer annular passage coaxially surrounds the core passage for directing a bypass gas flow. An exhaust mixer communicating with the core passage and outer annular passages has an upstream end, a downstream end and an annular wall extending therebetween. The annular wall defines a mixing plane at a trailing edge thereof at the downstream end at which the core and bypass gas flows are mixed. The annular wall has a radially inner surface and a radially outer surface. A plurality of protrusions are disposed on and extend laterally along and radially from at the radially inner surface and/or the radially outer surface. The plurality of protrusions are arranged one alongside another on the radially inner surface and/or the radially outer surface.

Abstract: An exhaust mixer assembly for a gas turbine engine includes a core passage extending along a central engine axis for directing a core gas flow. An outer annular passage coaxially surrounds the core passage for directing a bypass gas flow. An exhaust mixer communicating with the core passage and the outer annular passage has an upstream end, a downstream end and an annular wall extending therebetween. A plurality of protrusions extend axially from a downstream end of the annular wall to form a jagged trailing edge of the exhaust mixer. Each of the plurality of protrusions extend separately from the downstream end of the annular wall from the other of the plurality of protrusions. At least one of the plurality of protrusions is twisted about a protrusion axis extending through the at least one of the plurality of protrusions, the protrusion axis parallel to the central axis.

Abstract: An air intake for an aircraft propulsion system includes an air inlet duct, a core flow duct, a bypass flow duct, and a flow control device. The air inlet duct includes an intake inlet of the air intake. The core flow duct includes a core flow outlet. The core flow duct extends between and to the air inlet duct and the core flow outlet. The bypass flow duct includes a bypass flow outlet. The bypass flow duct extends between and to the air inlet duct and the bypass flow outlet. The bypass flow duct includes an interior surface forming and surrounding a bypass flow passage through the bypass flow duct. The flow control device is disposed on the interior surface. The flow control device is configured to variably control an area of a cross-sectional flow area of the bypass flow passage.