Abstract: A turbine shroud segment has a shroud body including a platform having forward and aft hooks extending from a radially outer surface of the platform for engagement with a shroud support structure of a turbine support case. A pin receiving hole is defined in the shroud body. An anti-rotation pin is engaged in the pin receiving hole. The anti-rotation pin projects outwardly from the pin receiving hole for engagement with a corresponding anti-rotation abutment on the shroud support structure.

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: A labyrinth seal assembly comprising: an outer seal having a radially inner cavity surface, and an axial cavity surface inward of the radially inner cavity surface; a floating seal received by the outer seal so as to be movable in a radial direction relative to the outer seal, the floating seal having a radially outer seal surface spaced inwardly from the radially inner cavity surface, a radially inner surface inward of the radially outer seal surface, and an axial seal surface bearing against the axial cavity surface; and an inner seal received by the floating seal, the inner seal having a radially outer surface spaced inwardly from the radially inner surface, a first one of the radially outer surface and the radially inner surface defining one tooth projecting toward a second one of the radially outer surface and the radially inner surface to inward of the axial seal surface.

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: A borescope plug assembly includes a housing, a plug, and a spring. The housing extends between a first housing end and a second housing end. The housing includes a plug aperture, an internal chamber, and a bayonet interface. The plug aperture is located at the second housing end. The internal chamber extends from the plug aperture toward the first housing end. The plug includes a shaft, a first sealing surface, and at least one tab. The shaft extends between a first shaft end and a second shaft end. The first sealing surface is disposed at the first shaft end. The at least one tab projects radially outward from the shaft. The plug extends through the plug aperture with the second shaft end and the at least one tab disposed within the internal chamber. The spring is positioned within the internal chamber between the first housing end and the at least one tab.

Abstract: A gas turbine intake has a swirl housing having a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the tangential inlet, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis, each vane having a twisted and flat body having a length extending from a leading end to a trailing end, the leading end being oriented mainly circumferentially and axially at the swirl path, the trailing end being oriented mainly axially and radially at the annular outlet, the twisted and flat body twisting between leading and trailing ends around the central axis, around a radial axis perpendicular to the central axis, and around a tangential axis perpendicular to both the central axis and the radial axis.

Abstract: A method of repair for a metallic part of an aircraft engine includes detecting a defect inside a bore of the metallic part wherein the defect represents a departure from an intended geometry of the bore, the bore having a diameter and defining a longitudinal axis, and wherein the defect is located within the bore at a depth of at least greater than one diameter along the longitudinal axis. The method also includes measuring a geometry of the defect, preparing a patch with a complementary geometry to fill the geometry of the defect, placing the patch in the bore with the complementary geometry of the patch seated against the geometry of the defect, directing a welding beam from outside the bore, through the bore and onto the patch to weld the patch to the bore, and removing a portion of the patch to provide the intended geometry for the bore.

Abstract: The compressor section can have a centrifugal impeller operable to rotate around an axis, the centrifugal impeller having blades, a compressor inlet oriented towards the front and axially relative the axis, a compressor outlet oriented radially outwardly relative the axis, a diffuser having a diffusion flow path, a diffuser inlet in fluid flow communication with the compressor outlet, a diffusion flow path between a rear wall and a front wall; a collector extending circumferentially around the axis, having a collector inlet in fluid communication with the diffuser outlet, and a collector outlet; and hollow structural members protruding rearwardly from the rear wall, the hollow structural members being circumferentially interspaced from one another, each hollow structural member having a length extending radially along the rear wall and having an internal conduit extending radially inwardly along the length.

Abstract: An aircraft fuel system comprises a fuel manifold adapter interchangeably connectable between respective fuel manifolds and flow divider valves of different aircraft engine platforms. The adapter has a body having an input interface and an output interface. The input interface is rigidly connectable via a rigid supply line to the flow divider valve of a given one of the different engine platforms so as to locate the output interface in axial alignment with a nozzle-input interface of the fuel manifold of the given one of the engine platforms. The output interface is axially slidably engaged with an upstream-tube end of a linear transfer tube having a downstream-tube end slidably engaged with the nozzle-input interface. The output interface of the adapter and the nozzle-input interface are slidably engaged and thermally decoupled via the linear transfer tube.

Abstract: The damper segment can be assembled between adjacent segments of a pressurized gas pipe of an aircraft engine. The damper segment can have a proximal end, a distal end, a rigid tube at the proximal end, a damper tube extending between the rigid tube and the distal end, the damper tube being made of a metal mesh, a proximal catch structurally connecting a proximal end of the damper tube to the rigid tube, and a distal catch structurally connected between a distal end of the damper tube and the distal end, the damper tube having an unsupported length extending between the distal catch and the proximal catch, the rigid tube having a liner portion projecting into the distal segment, the liner portion extending internally relative the damper tube.

Abstract: A fuel manifold adapter for a fuel system of an aircraft engine, the fuel manifold adapter comprising: a body having a body-output interface defining a downstream end of a body passage including a body bore about a bore axis, the body-output interface movably and fluidly connectable to a first component of the fuel system mounted to a first mounting point of the engine, and a body-input interface defining an upstream end of the body passage, the body-input interface rigidly and fluidly connectable to a second component of the fuel system mounted to a second mounting point of the engine, and a transfer tube having an upstream-tube end slidably engaged with the body along the bore axis via the body bore, the transfer tube having a downstream-tube end opposite the upstream-tube end slidably engageable along the bore axis with the first component, the downstream-tube end defining a downstream end of the fuel manifold adapter relative to fuel flow through the fuel manifold adapter.

Abstract: A probe shielding arrangement comprises a sleeve having a radially inner end mounted to a turbine housing and a radially outer end floatingly received in a probe boss on an exhaust case. The sleeve circumscribes an annular cavity around the probe. The annular cavity is sealed at opposed ends thereof to form a dead air cavity around the probe for insulation purposes.

Abstract: A probe shielding arrangement comprises a sleeve having a radially inner end mounted to a turbine housing and a radially outer end floatingly received in a probe boss on an exhaust case. The sleeve circumscribes an annular cavity around the probe. The annular cavity is sealed at opposed ends thereof to form a dead air cavity around the probe for insulation purposes.

Abstract: The gas turbine intake can have a swirl housing having an inlet portion fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the inlet portion to a circumferential outlet, the circumferential outlet fluidly connected back into the inlet portion, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis and located radially inwardly from the swirl path relative the central axis, the swirl path being free of the vanes.

Abstract: A gas turbine intake has a swirl housing having a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the tangential inlet, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis, each vane having a twisted and flat body having a length extending from a leading end to a trailing end, the leading end being oriented mainly circumferentially and axially at the swirl path, the trailing end being oriented mainly axially and radially at the annular outlet, the twisted and flat body twisting between leading and trailing ends around the central axis, around a radial axis perpendicular to the central axis, and around a tangential axis perpendicular to both the central axis and the radial axis.

Abstract: The gas turbine intake can have a swirl housing having an inlet portion fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending circumferentially around the central axis from the inlet portion to a circumferential outlet, the circumferential outlet fluidly connected back into the inlet portion, and vanes located in the swirl housing, the vanes circumferentially interspaced from one another relative the central axis and located radially inwardly from the swirl path relative the central axis, the swirl path being free of the vanes.

Abstract: An aircraft has an exhaust duct assembly disposed in a low-pressure turbine portion of a gas turbine engine to discharge exhaust gas to the outside and includes a duct having an inlet through which an exhaust gas is introduced, an exhaust portion through which the exhaust gas is exhausted, and a mounting bore formed therethrough; a duct housing in which the duct is disposed therein, and including a mounting hole into which the exhaust portion is inserted; a shaft housing disposed inside the mounting bore, inserted into the duct housing so that one end thereof faces the inlet of the duct, and having a driving axis of the low-pressure turbine portion disposed therein; and a sealing portion having a ring shape, being coupled to one end of the shaft housing, and maintaining airtightness between the driving axis and the shaft housing.

Abstract: An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a seal carrier, a seal land, a seal ring, a plate and a fastener. The seal carrier has an annular groove and extends between a first side and a second side. The seal land is opposite the annular groove. The seal ring seals a gap between the seal carrier and the seal land. The seal ring is seated in the annular groove. The plate is at the second side of the seal carrier. The fastener includes a head and an elongated member connected to the head. The head is at the first side of the seal carrier. The elongated member projects out from the head through the seal carrier, the seal ring and the plate. The elongated member is bonded to the plate.

Abstract: The gas turbine intake can have a swirl housing assembly with a tangential inlet fluidly connecting an exhaust conduit, an annular outlet defined around a central axis and fluidly connecting a turbine gas path, a swirl path extending around the central axis and between the tangential inlet and the annular outlet, the swirl housing assembly having a proximal portion defining a first portion of the swirl path, a distal portion defining a second portion of the swirl path, vanes located in the swirl housing assembly, the vanes circumferentially interspaced from one another relative the central axis and extending between the proximal portion and the distal portion, the proximal portion fastened to the distal portion via a plurality of fasteners, a gasket sandwiched between the proximal portion and the distal portion by the plurality of fasteners, the gasket extending in a radial plane relative the central axis.

Abstract: An apparatus is provided for a gas turbine engine. This apparatus includes a gas turbine engine case extending axially along and circumferentially around an axis. The gas turbine engine case includes a sheet of metal wrapped multiple times around the axis to form a containment structure having a multi-layered configuration. The containment structure is configured to contain at least one of a blade or a blade fragment from a bladed rotor of the gas turbine engine within the at least one of a plurality of sections.