Abstract: An engine carcass stiffener for a turboprop engine is provided. The engine carcass stiffener includes a first longitudinal reinforcement member, a proximal end of the first longitudinal reinforcement member is coupled to a first end of a proximal flange; a second longitudinal reinforcement member, a proximal end of the second longitudinal reinforcement member is coupled to a second end of the proximal flange; a first bushing holder, the first bushing holder is coupled to a first end of a distal flange; a second bushing holder, the second bushing holder is coupled to a second end of the distal flange; a first bolt, the first bolt is inserted through the first bushing holder and couples to a distal end of the first longitudinal reinforcement member; and a second bolt, the second bolt is inserted through the second bushing holder and couples to a distal end of the second longitudinal reinforcement member.

Abstract: An aircraft engine, has: a turbine including a turbine rotor rotatable about a central axis; a scroll case having an inlet fluidly connected to a source of combustion gases and an outlet fluidly connected to the turbine, and a conduit extending around the central axis from the inlet to the outlet, the conduit spiraling towards the central axis; a bearing housing extending around the central axis; an exhaust case disposed downstream of the turbine; and a turbine support case secured to the bearing housing and to the exhaust case, the turbine support case having spokes distributed around the central axis and extending along a direction having an axial component relative to the central axis, the spokes extending through the conduit of the scroll case and radially supported by the bearing housing.

Abstract: An engine carcass stiffener for a turboprop engine is provided. The engine carcass stiffener includes a first longitudinal reinforcement member, a proximal end of the first longitudinal reinforcement member is coupled to a first end of a proximal flange; a second longitudinal reinforcement member, a proximal end of the second longitudinal reinforcement member is coupled to a second end of the proximal flange; a first bushing holder, the first bushing holder is coupled to a first end of a distal flange; a second bushing holder, the second bushing holder is coupled to a second end of the distal flange; a first bolt, the first bolt is inserted through the first bushing holder and couples to a distal end of the first longitudinal reinforcement member; and a second bolt, the second bolt is inserted through the second bushing holder and couples to a distal end of the second longitudinal reinforcement member.

Abstract: Aircraft engines and methods of operation thereof are described. The aircraft engines include a fan section, a combustion section, a compressor section comprising an impeller, and a turbine section comprising a vane assembly and a containment cavity defined radially outward from the vane assembly. The fan section, the combustion section, the compressor section, and the turbine section are arranged along an engine axis and an engine shaft operably connects the turbine section to the fan section. Cooling air is supplied from the impeller to an inner diameter of the vane assembly, passed through at least one vane of the vane assembly, and into the containment cavity.

Abstract: An outer structural case of a turbine casing assembly, including: a plurality of threaded adjustment openings located about a periphery of the outer structural case, the plurality of threaded adjustment openings extending completely through the periphery of the outer structural case; and a plurality of threaded securement openings arranged orthogonally with respect to the plurality of adjustment openings.

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: An aircraft engine, has: a turbine including a turbine rotor rotatable about a central axis; a scroll case having an inlet fluidly connected to a source of combustion gases and an outlet fluidly connected to the turbine, and a conduit extending around the central axis from the inlet to the outlet, the conduit spiraling towards the central axis; a bearing housing extending around the central axis; an exhaust case disposed downstream of the turbine; and a turbine support case secured to the bearing housing and to the exhaust case, the turbine support case having spokes distributed around the central axis and extending along a direction having an axial component relative to the central axis, the spokes extending through the conduit of the scroll case and radially supported by the bearing housing.

Abstract: A gas turbine engine has a service tube assembly including an adjustable joint having a flange axially movable along an end fitting having a male/female connection with a mating part. The flange is threadably engaged with the end fitting to provide for the adjustment of the position of the flange along the end fitting.

Abstract: A mechanical dissuasion device includes a sleeve body forming a cylindrical body portion, a first tab, a second tab, and opposing fastener mounts. The cylindrical body portion extends circumferentially about a center axis. The cylindrical body portion extends circumferentially between and to a first circumferential end and a second circumferential end. The cylindrical body portion extends axially between and to a first axial end and a second axial end. The first tab is disposed at the first axial end. The first tab extends radially inward from the cylindrical body portion. The second tab is axially spaced from the first axial end and the second axial end. The second tab extends radially inward from the cylindrical body portion. The opposing fastener mount portions include a first fastener mount portion and a second fastener mount portion. The first fastener mount portion is disposed at the first circumferential end. The second fastener mount portion is disposed at the second circumferential end.

Abstract: A service tube assembly, has: a service tube having a threaded end portion and an outer shoulder adjacent to the threaded end portion, the outer shoulder having a frustoconical shoulder face converging towards the tube axis and away from the threaded end portion; a mating part threadingly engaged by the threaded end portion, the mating part having a seat face converging towards the service tube and towards the threaded end portion; a locking member extending from a first end to a second end, the first end including a notch receiving the outer shoulder, the notch circumscribed by a frustoconical notch face abutting the frustoconical shoulder face to create a frustoconical-to-frustoconical contact surface between the locking member and the outer shoulder, the second end in abutment against the seat face; and a mechanical fastener releasably holding the locking member in abutment against both of the outer shoulder and the mating part.

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 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 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 turbine assembly of an aircraft engine includes a cooling system for optimizing a tip clearance gap defined between an inner surface of a turbine housing and blade tips of the turbine blades. The cooling system includes a cooling airflow passage located radially outward from the turbine housing and being in heat-transfer communication with the turbine housing. The cooling airflow passage receives a flow of cooling air therethrough for cooling the turbine housing. A heat sink is disposed on the outer surface of the turbine housing within the cooling airflow passage, the heat sink including heat transfer elements projecting into the cooling airflow passage away from the outer surface of the turbine housing. The heat transfer elements are in convective heat transfer relationship with the flow of cooling air in the cooling airflow passage.

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: 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: 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: A service tube assembly comprises a service tube having a threaded end portion and a shoulder. The outer shoulder has a conical surface converging in a direction away from the threaded end portion. The threaded end portion is threadably engaged with a mating part. The mating part has a seat surface which cooperates with the shoulder of the service tube to form a conical seat. The assembly further includes a locking member having an outer conical surface complementary to the conical seat. A mechanical fastener is provided for releasably holding the locking member against the conical seat.

Abstract: A service tube assembly comprises a service tube having a threaded end portion and a shoulder. The outer shoulder has a conical surface converging in a direction away from the threaded end portion. The threaded end portion is threadably engaged with a mating part. The mating part has a seat surface which cooperates with the shoulder of the service tube to form a conical seat. The assembly further includes a locking member having an outer conical surface complementary to the conical seat. A mechanical fastener is provided for releasably holding the locking member against the conical seat.

Abstract: A turbine assembly in a turbine section of an aircraft engine includes a rotor with blades having blade tips, and a turbine housing radially surrounding the blades. A distance between an inner surface of the housing and the blade tips defines a tip clearance gap. A passive cooling system for optimizing the tip clearance gap includes a cooling airflow passage located radially outward from, and in heat-transfer with, the turbine housing. The cooling airflow passage has an inlet opening located upstream of the rotor and an exit opening located downstream of the rotor. The inlet opening provides air flow into the cooling airflow passage. The exit opening provides air flow communication between the cooling airflow passage and a main gaspath of the turbine section. A flow of cooling air through the cooling airflow passage is induced, to cool the housing.