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: A system and method for manufacturing a component is provided that includes a CNC machine tool, a correction module, and a system controller. The CMM module is controllable to determine a set of multi-axis coordinates surface points on the component. The correction module is in communication with the CMM module and stored reference inspection data. The system controller is in communication with the CNC machine tool, the correction module, and stored instructions. The instructions when executed cause the system controller to: a) control the CNC machine tool to modify a surface of the component; b) control the CMM module to determine multi-axis coordinates for surface points; c) determine surface position variances using the reference inspection data and the multi-axis coordinates; d) determine if surface position variances exceed a threshold; and e) create correction action instructions for controlling the CNC machine if surface position variances exceed the threshold.

Abstract: A fluid supply system and method is provided that includes a fluid pump, a pressure sensor, a pressure relief valve (PRV), and a fluid monitoring device. The fluid pump receives fluid from a first conduit and discharges fluid into a second conduit. The pressure sensor produces sensed fluid pressure signals. The PRV is in signal communication with the pressure sensor. The fluid monitoring device includes a control orifice in fluid communication with second and third conduits. The second conduit has a first inner diameter, the third conduit has a second inner diameter, and the control orifice has an orifice inner diameter, and the orifice inner diameter is less than the first and second inner diameters. The pressure sensor senses fluid pressure in the third conduit at a position in close proximity to the control orifice. The fluid monitoring device may be in a lead or a lag domain configuration.

Abstract: A measurement system for an aircraft gas turbine engine includes an instrumentation hub disposed about a rotational axis. The instrumentation hub includes at least one probe. The measurement system further includes a shield hub disposed about the rotational axis and positioned axially adjacent the instrumentation hub. The shield hub includes at least one shield configured to be radially aligned with the at least one probe of the instrumentation hub. The shield hub is rotatable about the rotational axis independent of the instrumentation hub. The at least one shield is axially translatable between a first axial position and a second axial position. The at least one shield in the first axial position is axially aligned with the at least one probe. The at least one shield in the second axial position is axially separated from the at least one probe.

Abstract: A tool assembly is provided for removing a component from a carrier. The component circumscribes and is mounted on the carrier. The tool assembly includes a tool head, an adaptor and an actuator. The tool head is configured to couple to the component. The adaptor extends longitudinally along a centerline between an adaptor first end and an adaptor second end. The adaptor is attachable to the tool head at the adaptor first end by a first quick coupler. The actuator includes a housing and a ram. The housing is attachable to the adaptor at the adaptor second end by a second quick coupler. The ram extends longitudinally along the centerline within the adaptor and the tool head to a ram distal end. The ram distal end is configured to engage the carrier.

Abstract: A structure is provided for a gas turbine engine. This gas turbine engine structure includes an engine case, an engine pylon and an engine line. The engine case includes a base, a mounting boss and a first support element. The base extends axially along and circumferentially about an axial centerline of the engine case. The mounting boss projects radially out from the base. The first support element projects radially out from the base and laterally out from the mounting boss. The first support element is configured as or otherwise includes a peripheral boss. The engine pylon is mounted to the mounting boss. The engine line is coupled to the peripheral boss.

Abstract: A diffuser pipe has a tubular body defining a pipe center axis extending therethrough. The tubular body includes a first portion extending in a generally radial direction from an inlet of the tubular body, a second portion extending in a generally axial direction and terminating at a pipe outlet, and a bend portion fluidly linking the first portion and the second portion. The tubular body has a length defined between the inlet and the pipe outlet. The tubular body has cross-sectional profiles defined in a plane normal to the pipe center axis. An area of the cross-sectional profile at the pipe outlet is at least 20% greater than an area of the cross-sectional profile at a point upstream from the pipe outlet a distance corresponding to 10% of the length of the tubular body.

Abstract: An assembly is provided for a turbine engine. This turbine engine assembly includes a static structure, a conduit and a bracket. The static structure includes a port. The conduit extends longitudinally through the port. The bracket couples the conduit to the static structure. The bracket includes a first base mount, a second bade mount, a conduit mount, a first damper and a second damper. The first base mount is attached to the static structure. The second base mount is attached to the static structure. The conduit mount is mechanically coupled with the conduit. The first damper is between the first base mount and the conduit mount. The second damper is between the second base mount and the conduit mount.

Abstract: An article has a nickel-based alloy substrate having, in weight percent: 5.4-7.4 Re; 4.1-5.9 Ru; 3.0-6.2 Cr; 3.0-10.0 Co; 0.5-3.8 Mo; 3.0-6.0 W; 4.6-8.6 Ta; 5.0-6.4 Al; 0.050-0.30 Hf; no more than 0.50 all other elements, if any, individually; and no more than 2.0 all other elements, if any, combined. A nickel-based coating is on the substrate and comprising, in weight percent: 6.0-10.0 Al; 4.0-15.0 Cr; 11.0-15.0 Co; 0.1-1.0 Hf; 0.1-1.0 Si; 0.1-1.0 Y; up to 1.0 Zr if any; up to 7.0 Ta if any; up to 6.0 W if any; no more than 1.0 all other elements, if any, individually; and no more than 4.0 all other elements, if any, combined.

Abstract: A combustor includes a combustor liner defining a combustion chamber and a fuel and air mixing body connected to the combustor liner to deliver mixed fuel and air into the combustion chamber. The mixing body includes an inner housing member centered on a center axis and an intermediate housing member. A mixing passage is defined between the inner and intermediate housing members. The mixing passage extends along a direction from an upstream end to a downstream end with a circumferential component, a component in an axially downstream direction, and a radially inward component with at least one air inlet into the mixing passage. A fuel supply extends into the mixing passage at a location downstream of the air inlet. The mixing passage extends downstream to supply fuel and air into the combustion chamber. A gas turbine engine is also disclosed.

Abstract: A propulsion system for an aircraft having a nacelle and a fuselage is provided. The nacelle has a gas flow path and a nacelle interior region. The system includes a compressor section, an intermittent IC engine, a turbine section, a fan, and an IC cooling system. A first electric motor powers the compressor section. The compressor section produces a flow of elevated pressure compressor air. The intermittent IC engine selectively intakes the compressor air flow and produces an exhaust gas flow. The turbine section powered by exhaust gas in turn powers a first electric generator. The fan is driven by a second electric motor. The IC engine cooling system has a heat exchanger disposed within the gas flow path, coolant, coolant piping, and a pump. The heat exchanger is disposed in the nacelle.

Abstract: A formation method is provided. During this formation method, a metallic substrate is provided. A coating is deposited onto the metallic substrate using a suspension plasma spray process. The coating is formed from or otherwise includes copper oxide.

Abstract: An assembly is provided for a gas turbine engine. This assembly includes a rotating structure, a stationary structure and a bushing. The rotating structure extends axially along and is rotatable about a centerline. The stationary structure extends circumferentially about the rotating structure. The stationary structure is configured from or otherwise includes stationary structure material with a coefficient of thermal expansion between 10 ?in/in-° F. and 15 ?in/in-° F. The bushing is radially between the rotating structure and the stationary structure. The bushing includes a mount and a bearing within the mount. The mount is configured from or otherwise includes mount material with a coefficient of thermal expansion between 9 ?in/in-° F. and 10 ?in/in-° F. The mount material contacts the stationary structure material. The bearing is configured from or otherwise includes bearing material, where the bearing material is engaged with and rotatably supports the rotating structure.

Abstract: An impeller shroud assembly for a gas turbine engine includes an annular impeller shroud disposed about an axial centerline. The impeller shroud includes a shroud inducer portion and a shroud exducer portion disposed radially outward of the shroud inducer portion and extending to an outer radial end of the impeller shroud. The shroud inducer portion and the shroud exducer portion defining an impeller-facing surface of the impeller shroud. The impeller shroud has a pivot point defined between the shroud inducer portion and the shroud exducer portion. The impeller shroud assembly further includes a clearance control device connected to the shroud exducer portion of the impeller shroud proximate the outer radial end. The clearance control device is configured to pivot the shroud exducer portion of the impeller shroud about the pivot point between a first axial position and a second axial position.

Abstract: A lubrication system is provided for a turbine engine. A lubricant source includes a source outlet. Feed circuits are fluidly coupled with the source outlet in parallel. The feed circuits include a first feed circuit and a second feed circuit. The second feed circuit includes a pump with a pump inlet and a pump outlet. A bypass circuit is fluidly coupled with the pump inlet and the pump outlet. A bleed circuit is fluidly coupled with the first feed circuit. A flow regulator is configured to regulate flow through the bypass circuit during a first mode of operation and a second mode of operation. The flow regulator is configured to close the bleed circuit during the second mode of operation. A sensor system is configured to monitor fluid flow directed to the first feed circuit and/or the second feed circuit.

Abstract: A gas turbine engine includes an intake device. The intake device includes a snorkel and a filter case. The snorkel includes a tubular body and an inlet aperture. The tubular body extends between a closed end and an open end opposite the closed end. The inlet aperture is formed through the tubular body proximate the closed end. The tubular body forms a first portion of a gas flow path for a bleed gas from the inlet aperture to the open end. The filter case is connected to the tubular body. The filter case extends between a first end and a second end. The filter case includes a sidewall extending from the first end to the second end. The sidewall surrounds a filter cavity. The filter case is configured to receive the bleed gas from the open end of the tubular body. The filter case and the snorkel form a unitary component.

Abstract: A fixture assembly includes a support stand, a support platform and a platform adaptor. The support platform is rotatably coupled to the support stand and rotatable about an axis. The support platform extends circumferentially about a centerline angularly offset from the axis. The support platform extends longitudinally along the centerline to a platform top surface. The platform adaptor is disposed on the platform top surface and attached to the support platform. The platform adaptor includes a case receptacle formed by an adaptor top surface and an adaptor shoulder surface. The platform adaptor extends longitudinally along the centerline to the adaptor top surface where the adaptor top surface is configured to longitudinally engage and support the case. The platform adaptor extends radially relative to the centerline to the adaptor shoulder surface where the adaptor shoulder surface is configured to radially engage and position the case on the adaptor top surface.

Abstract: A method is provided for disassembling a gas turbine engine. The gas turbine engine includes a compressor section, a combustor section, a turbine section, a static structure and a bypass duct. The static structure houses and supports the compressor section, the combustor section and the turbine section. The static structure includes a turbine support structure. The bypass duct includes an inner duct wall, an outer duct wall and a bypass flowpath formed radially between the inner duct wall and the outer duct wall. The outer duct wall extends axially along the static structure and overlaps the turbine support structure. During the method, the turbine support structure is removed from the gas turbine engine while the outer duct wall remains installed.

Abstract: A method is provided for testing an annular seal within a gas turbine engine. During this method, a vacuum is applied to a first volume through a conduit. The annular seal is between the first volume and a second volume. A vacuum pressure is measured within the conduit while the vacuum is applied. The measured vacuum pressure is compared to a threshold vacuum pressure. A difference between the measured vacuum pressure and the threshold vacuum pressure is indicative of leakage across the annular seal from the second volume to the first volume.

Abstract: An aircraft system is provided that includes a first propulsor rotor, a second propulsor rotor, a drivetrain and an intermittent combustion engine. The first propulsor rotor is rotatable about a first propulsor axis. The second propulsor rotor is rotatable about a second propulsor axis. The drivetrain includes a drive structure and a transmission. The drive structure is rotatable about a drive axis that is angularly offset from the first propulsor axis and the second propulsor axis. An output of the transmission is coupled to the first propulsor rotor and the second propulsor rotor through the drive structure. The intermittent combustion engine is configured to drive rotation of the first propulsor rotor and the second propulsor rotor through the drivetrain.