Abstract: A reduction gearbox is provided that includes a planetary gear arrangement disposed in a casing. The planetary gear arrangement includes a sun gear, planet gear assemblies, and a fore and aft ring gears. The sun gear is configured in communication with a cantilevered end of an input. Each planet gear assembly has a main gear meshed with the sun gear, a fore lateral gear and an aft lateral gear disposed on opposite sides of the main gear and rotating therewith. The main gear has a main gear pitch diameter and the fore and aft lateral gears having a lateral gear pitch diameter. The fore ring gear is meshed with the fore lateral gears. The aft ring gear is meshed with the aft lateral gears. The aft ring gear is independent of the fore ring gear. The planet gear assemblies are in communication with an output.

Abstract: A combustor includes a liner defining a combustion chamber and receiving a fuel and air mixing body. The mixing body has a central fuel supply, and radial distribution passages communicating fuel from the central fuel supply radially outwardly relative to a central axis of the central fuel supply and to mixing passages. The radial distribution passages have injection ports in the mixing passages. The mixing passages extend from a rear face of the mixing body to an inner face facing into the combustion chamber. Air inlets in the mixing body communicate air into the mixing passages, and there is cellular material in the mixing passages at a location at which the fuel is injected into the mixing passages. A gas turbine engine is also disclosed.

Abstract: A propulsion system for an aircraft includes an engine, a plurality of sensors, and a controller. The engine includes an air intake. The plurality of sensors includes an outside air temperature (OAT) sensor and an air intake sensor. The OAT sensor is disposed outside the engine and configured to measure an OAT. The air intake sensor is disposed inside the air intake and configured to measure an air intake temperature (T1). The controller is configured to: calculate an actual charge heating value (CHactual), compare the CHactual to a charge heating threshold value (CHthresh) to identify the CHactual is greater than or less than the CHthresh, identify a presence or an absence of an unusual engine performance condition for the engine, and identify a presence or an absence of a blockage condition of the air intake based on the CHactual greater than the CHthresh and the presence of the unusual engine performance condition.

Abstract: A method for qualifying a manufacturing process for a first component including a common geometric feature includes obtaining manufacturing data for the common geometric feature. The manufacturing data is associated with one or more qualified second components. The one or more qualified second components are different than the first component. Each of the one or more qualified second components include the common geometric feature. The method further includes modeling the manufacturing process for the common geometric feature of the one or more qualified second components using the manufacturing data, modeling the manufacturing process for the common geometric feature of the first component, obtaining manufacturing process parameters for the manufacturing process for the common geometric feature of the one or more qualified second components, and qualifying the manufacturing process for the common geometric feature of the first component.

Abstract: A fluid system for an aircraft engine can include a pumped fluid circulation circuit including a fluid tank, the fluid circulation circuit configured to circulate a fluid through at least a part of the aircraft engine and means, in the fluid circulation circuit, for selectively directing the fluid of the fluid circulation circuit as a function of a pressure of fluid in the fluid circulation circuit to one of: a part of the aircraft engine, and a fluid bypass in the fluid circulation circuit that is fluidly upstream of the part.

Abstract: An aircraft propulsion system and method of cooling the same are provided. The system includes a hydraulic pump, an intermittent IC engine, a hydraulic motor, an engine oil pump, and a cooling system. The intermittent IC engine drives the hydraulic pump. The hydraulic motor is powered by the hydraulic pump and drives a propulsor fan. The cooling system includes a first heat exchanger (AIR-EO HEX) and a second heat exchanger (EO-HF HEX). The AIR-EO HEX transfers heat between flows of engine oil and ambient air. The EO-HF HEX transfers heat between flows of engine oil and hydraulic fluid. The hydraulic pump provides motive force to cause the hydraulic oil to pass through the EO-HF HEX and back to the at least one hydraulic pump. The engine oil pump provides motive force to pass the engine oil to and through the AIR-EO HEX, the EO-HF HEX, and the intermittent IC engine.

Abstract: A strut includes a strut body extending in a radial direction and defining an airfoil shape in cross-section perpendicular to the radial direction. The airfoil shape includes a leading edge and a trailing edge. An extraction inlet is defined through an exterior surface of the strut body, in fluid communication with an internal conduit of the strut body. An injection outlet is defined through the exterior surface of the strut body, in fluid communication with the internal conduit for fluid communication through the internal conduit from the extraction inlet to the injection outlet.

Abstract: The rotor control system includes a main control valve having an inlet for receiving liquid and an outlet for issuing liquid to a rotor pitch change actuator. The rotor control unit is configured to control flow of liquid from the inlet to the outlet to modify pitch angle of rotor blades. A feathering system with a first conduit connected in fluid communication with the outlet of the main control valve, a second conduit for fluid communication with the rotor pitch change actuator, and a drain conduit. The feathering system has a normal operation mode for supplying liquid from the main control valve to the rotor pitch change actuator by allowing flow through the feathering system from the first conduit to the second conduit. The feathering system has a feathering mode for controlling drainage from the rotor pitch change actuator to the drain conduit.

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.

Abstract: An apparatus is provided for a gas turbine engine. This apparatus includes a lubricant reservoir, a heat exchanger, a lubricant circuit and a compressed air circuit. The lubricant reservoir includes an internal cavity for storing lubricant. The heat exchanger includes a heat exchange conduit and a compressed air passage. The heat exchange conduit is arranged in the internal cavity. The compressed air passage extends longitudinally through the heat exchange conduit. The lubricant circuit is fluidly coupled with the internal cavity. The compressed air circuit is fluidly coupled with the compressed air passage.

Abstract: An integrally bladed rotor (IBR) for a gas turbine engine is provided. The IBR includes a hub, rotor blades that include a test blade, and at least one heating element. Each rotor blade has an airfoil with leading and trailing edges, suction side and pressure side surfaces, and a base end. The airfoil of the test blade includes at least one slot defining a void in the airfoil. The slot extends a lengthwise distance into the airfoil along a direction generally between the leading and trailing edges of the airfoil and terminates at a slot end surface. The airfoil includes at least one internal cavity extending lengthwise from the slot end surface. The heating element is disposed in the internal cavity and selectively produces thermal energy sufficient to heat the airfoil material proximate the internal cavity to a temperature at which the airfoil mechanical strength properties are decreased.

Abstract: A method is provided for disassembling a rotor of a gas turbine engine. During this method, the rotor is provided which includes a rotor disk and a plurality of rotor blades arranged circumferentially about an axis. The rotor blades include a plurality of airfoils and a plurality of attachments that mount the rotor blades to the rotor disk. Each of the rotor blades includes a respective one of the airfoils and a respective one of the attachments. A press is arranged against the rotor. The press axially engages each of the rotor blades. The press moves axially along the axis to simultaneously push the rotor blades and remove the attachments from a plurality of slots in the rotor disk.

Abstract: An air intake for an aircraft propulsion system includes an air inlet duct and a flow control device. The air inlet duct includes an interior surface and an intake inlet. The interior surface extends from the intake inlet. The interior surface forms and surrounds an inlet flow passage through the air inlet duct. The intake inlet includes a top side, a bottom side, a first lateral side, and a second lateral side. Each of the first lateral side and the second lateral side extend between and to the top side and the bottom side. The flow control device is disposed inside the air inlet duct at the first lateral side. The flow control device is configured to form an asymmetrical shape of the intake inlet at the first lateral side relative to the second lateral side.

Abstract: A method is provided for operating an aircraft system. During this method, an engine is operated using first fuel provided by a first fuel source. A fuel supply for the engine is switched between the first fuel source and a second fuel source, where the switching of the fuel supply includes shutting down the engine during aircraft flight. The engine is operated using second fuel provided by the second fuel source.

Abstract: A rotor for an aircraft rotary engine includes a rotor body. The rotor body extends about an axial centerline. The rotor body includes an outer body portion, an annular inner body portion, and ribs. The outer body portion forms a plurality of sides and a plurality of apex portions of the rotor. The annular inner body portion is disposed radially inward of the outer body portion. The ribs extend radially between and connect the outer body portion and the annular inner body portion. The rotor body forms at least a first lubrication passage within the inner body portion and within a first rib of the ribs. The first lubrication passage includes at least one passage inlet and at least one passage outlet. The at least one passage inlet is disposed at the annular inner body portion.

Abstract: An air intake for an aircraft propulsion system includes an air inlet duct, a core flow duct, a bypass flow duct, a splitter, and a flow control device. The air inlet duct includes an intake inlet and a gas path floor. The core flow duct includes a core flow outlet. The bypass flow duct includes a bypass flow outlet. The bypass flow duct includes the gas path floor. The splitter separates the core flow duct and the bypass flow duct. The flow control device is disposed on a portion of the gas path floor. The flow control device is configured to be selectively positioned to control an air flow path for air flowing through the air inlet duct, the core flow duct, and the bypass flow duct.

Abstract: An assembly is provided for a gas turbine engine. This engine assembly includes a bladed rotor rotatable about an axis, and an engine case. The engine case includes an outer duct wall, a first circumferential stiffener, a second circumferential stiffener and a plurality of axial stiffeners. The outer duct wall forms a shroud around the bladed rotor. The first circumferential stiffener extends circumferentially about the outer duct wall. The second circumferential stiffener extends circumferentially about the outer duct wall. The axial stiffeners are arranged circumferentially about the outer duct wall. Each of the axial stiffeners extends axially between the first circumferential stiffener and the second circumferential stiffener.

Abstract: An integrally bladed rotor (IBR) for a gas turbine engine and method is provided. The IBR is configured for use in blade off testing and includes a hub, a plurality of rotor blades, a central passage, and first and second lateral cavities. The hub has forward and aft ends and a circumferentially extending exterior surface. The central passage is disposed in the hub radially below a test rotor blade, extending along a path between an inlet at or forward of the test blade leading edge and an outlet at or aft of the test blade trailing edge. The first and second lateral cavities are disposed in the hub, extending generally parallel to the central passage path, on opposite circumferential sides. The first lateral cavity is disposed a distance (MSD1) from the central passage and the second lateral cavity is disposed a distance (MSD2) from the central passage.

Abstract: A fuel assembly for a gas turbine engine includes a fuel supply tube, a fuel port, a fuel manifold, and a fuel manifold adapter. The fuel supply tube is configured to convey a fuel. The fuel port is fluidly coupled to the fuel supply tube and configured to receive the fuel from the fuel supply tube. The fuel manifold includes a fuel inlet and a plurality of fuel outlets. The fuel inlet is fluidly coupled to the fuel port and configured to receive the fuel from the fuel port. The fuel manifold adapter includes a first mount portion and a second mount portion. The first mount portion is connected to the fuel port. The first mount portion is moveable relative to the second mount portion.

Abstract: During an inspection method, an inspection scope is inserted into an interior of an apparatus of a gas turbine engine. An image of the interior of the apparatus is captured to provide/output image data. A location of the inspection scope within the interior of the apparatus when the image is captured is determined to provide location data, where the determining of the location includes comparing the image data to model data from a model of the apparatus. During the course of inspection, movement of the inspection scope is tracked within the interior of the apparatus using the location data. The location data may be derived by feature recognition and/or motion tracking of live feed from a scope video output to features available in a field of view within duplicate CAD model.