Abstract: A casing for a gas turbine engine, including: a containment fabric layer wrapped around a structural shell; and a bi-material shell located around the containment fabric layer, the bi-material shell formed from a plurality of bi-material shell segments, each of the plurality of bi-material shell segments include an inner layer and an outer layer, the outer layer comprises a plurality of individual segments that form the outer layer and a plurality of embedded fuse lines defined by the plurality of individual segments of the outer layer, wherein an overall structure of the bi-material shell does not impede a rapid expansion of the containment fabric layer.

Abstract: A casing for a gas turbine engine, including: an inner surface portion; a pair of wall portions extending radially outward from the inner surface portion; and a containment ring located between the pair of wall portions, the containment ring including an inner ring, an intermediate ring, and an outer ring, the intermediate ring being located between the inner ring and the outer ring, wherein the containment outer ring is mechanically decoupled from the pair of wall portions.

Abstract: A system for detecting a shaft shear comprises a gas turbine engine having a turbine, a high inertia rotating component and a driveshaft. The turbine drives the high inertia rotating component using the driveshaft. A first speed probe is configured to detect a first speed of the driveshaft at a first location associated with the driveshaft. A second speed probe is configured to detect a second speed of the driveshaft at a second location associated with the driveshaft. Control circuitry is configured to compare the first speed and second speed, determine if the shaft shear has occurred responsive to the comparison and generate a control signal to the gas turbine engine responsive to the determination.

Abstract: A system for detecting shaft shear comprises a gas turbine engine including at least one turbine configured to drive a fan using a driveshaft. A plurality of frangible plugs is located downstream of the at least one turbine of the gas turbine engine. The plurality of frangible plugs is configured to provide an indication of shearing of the driveshaft responsive to breakage of at least a portion of the plurality of frangible plugs caused by impact of the at least one turbine with the portion of the plurality of frangible plugs.

Abstract: A propulsion system for an aircraft includes an engine assembly, an interburner, a turbocompressor, and a flow control assembly. The engine assembly includes an engine. The engine includes an air inlet, an exhaust outlet, and an engine output shaft. The interburner includes an interburner inlet and an interburner outlet. The turbocompressor assembly includes at least one turbocompressor. The at least one turbocompressor includes a turbine and a compressor. The turbine includes a turbine inlet and a turbine outlet. The compressor includes a compressor inlet and a compressor outlet. The flow control assembly includes an engine exhaust diverter subassembly and at least one compressor outlet diverter subassembly. The engine exhaust diverter subassembly is connected to exhaust outlet, the interburner inlet, and the turbine inlet. The at least one compressor outlet diverter subassembly is connected to the compressor outlet, the air inlet, and the interburner inlet.

Abstract: An aircraft powerplant assembly includes a compressor section, a combustor section, a turbine section, a flowpath and a rotating structure. The flowpath extends through the compressor section, the combustor section and the turbine section. The rotating structure includes a balancing device and a bladed rotor disposed in the turbine section or the compressor section. The balancing device is configured to rotationally balance the rotating structure about a rotational axis of the rotating structure. The balancing device includes a mounting platform and a plurality of balancing masses. The mounting platform includes a plurality of mounting apertures. The balancing masses are arranged circumferentially about the axis. Each of the balancing masses includes a shank and a head. The shank of each of the balancing masses is threaded into a respective one of the mounting apertures. The head of each of the balancing masses is disposed radially inboard of the mounting platform.

Abstract: A casing for a gas turbine engine, including: an inner ring portion; an outer ring portion; and a plurality of blunting plates arranged circumferentially between the inner ring portion and the outer ring portion, wherein the plurality of blunting plates are positioned to overlap each other.

Abstract: A bearing arrangement of a gas turbine engine includes a first bearing supportive of a shaft, a second bearing axially offset from the first bearing, a first bearing support extending from a bearing housing to the first bearing to support the first bearing and a second bearing support extending to and supportive of the second bearing. One or more radial bearing supports include a plurality of radial spokes extending between the bearing housing and the second bearing. The plurality of radial spokes are configured to buckle under compression when radial loads at the second bearing exceed a predetermined threshold. At least one destabilizer is secured to the second bearing and is configured to contact at least one radial spoke of the plurality of radial spokes upon failure of the first bearing support to reduce the predetermined threshold sufficiently to initiate buckling of the plurality of radial spokes.

Abstract: A bearing arrangement of a gas turbine engine includes a first bearing supportive of a shaft, and a second bearing axially offset from the first bearing. A first bearing support extends from a bearing housing to the first bearing, and a second bearing support extends from a bearing housing to the second bearing. One or more radial bearing supports including a plurality of radial spokes extend between the bearing housing and the second bearing. The spokes are configured to buckle under compression when radial loads at the second bearing exceed a predetermined force threshold. An axially-extending spoke foot extends from at least one radial spoke. The axially-extending foot is configured to contact the second bearing in the event of a tilt of the second bearing beyond a predetermined angle threshold, reducing the buckling force of the plurality of radial spokes below a second predetermined force threshold.

Abstract: A computer-readable model of target speckle pattern and a method for constructing the computer-readable model of target speckle pattern are described. The computer-readable model is constructed with computer-stimulated digital image correlation (DIC) unit that includes a camera and a three-dimensional object model. The three-dimensional object model is positioned within a view frustum of the camera and facing the camera. A nominal speckle pattern is generated based on the parameters of the camera. The camera projects the nominal speckle pattern onto the three-dimensional object model so that a target speckle pattern can be constructed. The target speckle pattern is then saved in a non-transitory computer-readable memory of the computer-readable model. The computer-readable model may be executed by a controller of a manufacturing system, such as a CNC machining system or a 3D printing system, to manufacture instances of the stencil.

Abstract: A rotating structure for an aircraft powerplant includes a balancing device and a bladed rotor. The balancing device is configured to rotationally balance the rotating structure about a rotational axis. The balancing device includes a housing and a plurality of balancing masses. The housing includes a first sidewall, a second sidewall, an outer endwall, an annular channel and an annular slot. The outer endwall circumscribes the annular channel. The annular channel extends axially within the housing between the first sidewall and the second sidewall. The annular slot extends radially through the outer endwall to the annular channel. Each of the balancing masses includes a mass body and an engagement feature. The mass body is captured within the annular channel. The engagement feature is configured to be engaged by a tool from outside of the housing to adjust a circumferential position of the mass body about the rotational axis.

Abstract: Systems for actuating a plurality of guide vanes arranged about an engine axis include a unison ring is operably connected to each of the guide vanes to impart rotation to the guide vanes about a respective guide vane axis. A guide vane actuator assembly includes an actuator, an actuating rod, a boss rod connected to the actuating rod with the actuating rod connected to a slotted end of the boss rod, a jointed boss member defining a through hole through a housing and supporting the boss rod, and a connecting rod operably connecting the boss rod to a bracket, with the bracket connected to the unison ring. The actuator is configured to apply an actuating force to the actuating rod to apply the actuating force to the unison ring and control the rotation of the guide vanes about the respective guide vane axes.

Abstract: A propulsion system for an aircraft includes an engine assembly and a turbocompressor. The engine assembly includes an engine and an interburner. The engine includes an engine output shaft. The engine is configured to drive rotation of a propulsor with the engine output shaft. The interburner is configured to mix and burn an exhaust gas from the engine with fuel to form a combustion gas. The turbocompressor includes a turbine and a compressor. The turbine and the compressor form a rotational assembly. The rotational assembly includes a shaft, a bladed turbine rotor of the turbine, and a bladed compressor rotor of the compressor. The turbine is connected in fluid communication with the interburner to receive the combustion gas. The compressor is connected in fluid communication with the engine to direct a compressed air to the engine. The rotational assembly is mechanically independent of the engine output shaft.

Abstract: A propulsion system for an aircraft includes an air intake, an engine assembly, a turbocompressor, an electrical assembly, and at least one propulsor. The engine assembly includes an engine. The engine includes an air inlet, an exhaust outlet, and an engine output shaft. The turbocompressor assembly includes a turbocompressor. The turbocompressor includes a turbine and a compressor. The turbine and the compressor form a rotational assembly. The rotational assembly includes a shaft, a bladed turbine rotor of the turbine, and a bladed compressor rotor of the compressor. The shaft interconnects the bladed turbine rotor and the bladed compressor rotor. The turbine includes a turbine inlet and a turbine outlet. The turbine inlet is connected in fluid communication with the exhaust outlet. The compressor includes a compressor inlet and a compressor outlet. The compressor inlet is connected in fluid communication with the air intake. The rotational assembly is mechanically independent of the engine output shaft.

Abstract: An assembly is provided for a rotorcraft. This rotorcraft assembly includes a fuselage, a tail boom, an air system and a cooling system. The tail boom projects longitudinally along a centerline out from the fuselage. The air system includes an air flowpath. The air flowpath passes from the fuselage into the tail boom and extends longitudinally within the tail boom. The cooling system includes a first heat exchanger and a coolant flowpath extending through the first heat exchanger. The first heat exchanger is arranged with the tail boom. The first heat exchanger is configured to transfer heat energy between system air flowing in the air flowpath and coolant flowing in the coolant flowpath.

Abstract: An assembly for an aircraft includes a first engine case, a bracket and fastener assemblies. The first engine case includes a first sidewall and a first flange projecting radially out from the first sidewall. The bracket is abutted axially against the first flange. The fastener assemblies attach the bracket to the first flange. The fastener assemblies include a first fastener assembly with a first bolt, a first nut and a retainer. The first bolt includes a first bolt head and a first bolt shank that projects axially out from the first bolt head and axially through the first flange and the bracket. The first nut is threaded onto the first bolt shank with the first flange and the bracket is axially between the first bolt head and the first nut. The retainer is attached to the first bolt shank with the first nut axially between the retainer and the bracket.

Abstract: A rotating structure for an aircraft powerplant includes a shaft, a retainer, a balancing device and a bladed rotor. The shaft includes a bore and a shoulder axially along the bore. The retainer is disposed in the bore and mounted to the shaft. The balancing device is configured to rotationally balance the rotating structure about the axis. The balancing device is disposed in the bore and axially retained between the shoulder and the retainer. The balancing device includes a balancing mass with a plurality of mass teeth arranged circumferentially about the axis. During a first condition, the balancing mass is rotationally repositionable about the axis relative to the shaft while the balancing device remains in the bore and axially retained between the shoulder and the retainer. During a second condition, the mass teeth are meshed with a plurality of shaft teeth to rotationally fixed the balancing mass to the shaft.

Abstract: A bearing arrangement of a gas turbine engine includes a first bearing and a second bearing axially offset from the first bearing. A first bearing support extends from a bearing housing to the first bearing to support the first bearing, and a second bearing support extends from the bearing housing to the second bearing to support the second bearing. One or more radial bearing supports including a plurality of radial spokes extend between the bearing housing and the second bearing. Each spoke extends from a spoke base to a spoke tip. A raised pedestal is positioned at a contact point of each spoke tip, and at least one destabilizer is secured to the second bearing and is configured to contact at least one radial spoke when the predetermined threshold is exceeded to initiate sliding of the plurality of radial spokes from their respective raised pedestal.

Abstract: A bearing arrangement of a gas turbine engine includes a bearing supportive of a shaft, a bearing support extending from the bearing housing to the bearing to support the bearing, and one or more radial bearing supports including a plurality of radial spokes extending between the bearing housing and the bearing. The plurality of radial spokes are configured to buckle under compression when radial loads at the bearing exceed a predetermined threshold.

Abstract: A bearing arrangement of a gas turbine engine includes a first bearing and a second bearing axially offset from the first bearing. A first bearing support extends from a bearing housing to the first bearing to support the first bearing, and a second bearing support extends from the bearing housing to the second bearing to support the second bearing. One or more radial bearing supports including a plurality of radial spokes extend between the bearing housing and the second bearing. Each spoke extends from a spoke base to a spoke tip. A raised pedestal is positioned at a contact point of each spoke tip, and at least one destabilizer is secured to the second bearing and is configured to contact at least one radial spoke when the predetermined threshold is exceeded to initiate sliding of the plurality of radial spokes from their respective raised pedestal.