Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.

Abstract: A method for operating a hybrid electric propulsion system of an aircraft, the hybrid electric propulsion system comprising a turbomachine, an electric machine coupled to the turbomachine, and a propulsor coupled to the turbomachine, the method comprising: operating the turbomachine to drive the propulsor; receiving data indicative of a failure condition of the hybrid electric propulsion system; and extracting power from the turbomachine using the electric machine to slow down one or more rotating components of the turbomachine in response to receiving the data indicative of the failure condition.

Abstract: A method for operating a hybrid electric propulsion system of an aircraft, the hybrid electric propulsion system comprising a turbomachine, an electric machine coupled to the turbomachine, and a propulsor coupled to the turbomachine, the method comprising: operating the turbomachine to drive the propulsor; receiving data indicative of a failure condition of the hybrid electric propulsion system; and extracting power from the turbomachine using the electric machine to slow down one or more rotating components of the turbomachine in response to receiving the data indicative of the failure condition.

Abstract: A method and icing effects mitigation system are provided. The icing effects mitigation system includes a fluid duct configured to channel a first flow of fluid through the fluid duct from a duct opening to a rotatable member at least partially positioned within the fluid duct. The rotatable member includes a radially inner rotatable portion and a radially outer rotatable portion. The icing effects mitigation system also includes a duct member extending through the fluid duct in a direction approximately orthogonal to a direction of the first flow of fluid. The duct member is configured to channel a second flow of a second fluid therethrough that causes ice accreted on the duct member to shed on a trajectory that impacts the rotatable member at the radially inner portion.

Abstract: A method and icing effects mitigation system are provided. The icing effects mitigation system includes a fluid duct configured to channel a first flow of fluid through the fluid duct from a duct opening to a rotatable member at least partially positioned within the fluid duct. The rotatable member includes a radially inner rotatable portion and a radially outer rotatable portion. The icing effects mitigation system also includes a duct member extending through the fluid duct in a direction approximately orthogonal to a direction of the first flow of fluid. The duct member is configured to channel a second flow of a second fluid therethrough that causes ice accreted on the duct member to shed on a trajectory that impacts the rotatable member at the radially inner portion.

Abstract: Embodiments of hot streak alignment for gas turbine durability include structures and methods to align hot streaks with the leading edges of aligned first stage nozzle vanes in order to improve mixing of the hot streaks with cooling air at a stator nozzle of a first stage turbine and reduce usage of cooling air at first stage non-aligned stator nozzle vanes disposed adjacent to the aligned stator vanes.

Abstract: Active clearance control ejector systems and methods are disclosed. An example system may include an air ejector; an entrance pipe arranged to supply compressor bleed air to the air ejector; an inlet duct arranged to supply fan bypass air to the air ejector; and a supply pipe arranged to receive ejector outlet air from the air ejector and supply the ejector outlet air to an active clearance control system. The ejector outlet air may include a mixture of the compressor bleed air and the fan bypass air. The air ejector may include a venturi arranged to conduct the compressor bleed air therethrough, thereby drawing the fan bypass air into the air ejector.

Abstract: A connection between an annular sleeve on a gas turbine engine first rotor component includes an annular tapered conical slot extending axially inwardly into the sleeve from an annular opening and tapers axially into the sleeve from the opening. An annular rim of a second rotor component is received within the conical slot and includes a rim inner conical surface mating with and contacting a sleeve inner conical surface in part bounding the conical slot. A sleeve cylindrical outer surface in part bounds the conical slot in the sleeve and a rim outer cylindrical surface on the annular rim contacts the sleeve cylindrical outer surface. An annular lip may extend radially inwardly from the sleeve and sleeve inner conical surface at the annular opening to the conical slot. The connection may be used between a first stage disk and an interstage seal.

Abstract: A method and system for a rotatable member of a turbine engine are provided. The rotatable member includes a substantially cylindrical shaft rotatable about a longitudinal axis, and a hub coupled to the cylindrical shaft through a conical shaft portion wherein the conical shaft portion includes a plurality of circumferentially-spaced air passages and wherein at least one of the plurality of air passages includes a non-circular cross section.

Abstract: A connection between an annular sleeve on a gas turbine engine first rotor component includes an annular tapered conical slot extending axially inwardly into the sleeve from an annular opening and tapers axially into the sleeve from the opening. An annular rim of a second rotor component is received within the conical slot and includes a rim inner conical surface mating with and contacting a sleeve inner conical surface in part bounding the conical slot. A sleeve cylindrical outer surface in part bounds the conical slot in the sleeve and a rim outer cylindrical surface on the annular rim contacts the sleeve cylindrical outer surface. An annular lip may extend radially inwardly from the sleeve and sleeve inner conical surface at the annular opening to the conical slot. The connection may be used between a first stage disk and an interstage seal.

Abstract: A method and system for a rotatable member of a turbine engine are provided. The rotatable member includes a substantially cylindrical shaft rotatable about a longitudinal axis, and a hub coupled to the cylindrical shaft through a conical shaft portion wherein the conical shaft portion includes a plurality of circumferentially-spaced air passages and wherein at least one of the plurality of air passages includes a non-circular cross section.

Abstract: A gas turbine engine disk that includes a centrally disposed disk hub having an integrally-formed, radially outwardly extending web terminating at an outer end. The disk hub has a radially-displaced annular hub surface exposed to high pressure, high temperature discharge gases during engine operation. The radially-displaced annular hub surface acts as an axial free surface mitigating the formation undesirable axial stress in the disk hub.

Abstract: A first stage turbine disk section for a gas turbine engine, including a centrally disposed disk hub having an integrally-formed, radially outwardly extending web terminating at an outer end. The disk hub has a convex, curved hub surface exposed to high pressure, high temperature compressor bleed and discharge gases during engine operation, and the convex, curved hub surface acts to reduce the axial constraint at the location of peak axial stress, thereby mitigating the formation of undesirable axial stress in the disk hub and allows separation of the peak axial and hoop stresses, thereby reducing the magnitude of the peak hub surface effective stress.

Abstract: A gas turbine section of a gas turbine engine, including a first stage turbine disk, a forward outer seal carried on the turbine disk and a forward shaft cooperating with the first stage turbine disk for transmitting torque generated by the turbine section to a compressor section of the engine. The forward shaft intersects the first stage turbine disk at a non-perpendicular angle in the area of the disk hub for directing sufficient hot air against the disk hub to balance the thermal response rate of the disk hub to the thermal expansion rate of the forward outer seal.

Abstract: A gas turbine engine blade root shim includes a dovetail shim portion with a dovetail shape and a longitudinally extending substantially flat base and distal first and second longitudinally spaced apart forward and aft ends. Transversely spaced apart first and second walls extend upwardly from the base which includes at least two longitudinally extending elongated base apertures. Each of the base apertures includes a main region and longitudinally spaced apart rounded end regions. Dog-bone-shaped base apertures or rounded end base apertures having semi-circular rounded end regions may be used. The shim is designed for a blade dovetail root with a root slot formed in a bottom of the dovetail root. The dovetail shim portion with the dovetail shape substantially conforms to at least a portion of a cross-sectional dovetail shape of the dovetail root through the root slot. The base is disposed within the root slot.

Abstract: A gas turbine engine blade root shim includes a dovetail shim portion with a dovetail shape and a longitudinally extending substantially flat base and distal first and second longitudinally spaced apart forward and aft ends. Transversely spaced apart first and second walls extend upwardly from the base which includes at least two longitudinally extending elongated base apertures. Each of the base apertures includes a main region and longitudinally spaced apart rounded end regions. Dog-bone-shaped base apertures or rounded end base apertures having semicircular rounded end regions may be used. The shim is designed for a blade dovetail root with a root slot formed in a bottom of the dovetail root. The dovetail shim portion with the dovetail shape substantially conforms to at least a portion of a cross-sectional dovetail shape of the dovetail root through the root slot. The base is disposed within the root slot.

Abstract: A method for assembling a rotating X-ray tube, the X-ray tube having a cathode for emitting electrons, and a rotor and a bearing assembly for facilitating rotation of an anode. The method includes using an interference fit assembly between the bearing assembly and the rotor to provide a joint having balance retention. The using interference fit assembly further includes selecting a rotor hub of the rotor having a coefficient of thermal expansion which results in a composite coefficient of thermal expansion for the rotor assembly which closely matches that of a shaft of the bearing assembly. An interference fit joint between a shaft extending from a bearing assembly and a rotor hub is also disclosed, wherein the joint is completed without using any mechanical fasteners or metallurgical bonding, yet can carry all of the operational loads placed on the joint.

Abstract: A method for assembling a rotating X-ray tube, the X-ray tube having a cathode for emitting electrons, and a rotor and a bearing assembly for facilitating rotation of an anode. The method includes using an interference fit assembly between the bearing assembly and the rotor to provide a joint having balance retention. The using interference fit assembly further includes selecting a rotor hub of the rotor having a coefficient of thermal expansion which results in a composite coefficient of thermal expansion for the rotor assembly which closely matches that of a shaft of the bearing assembly. The shaft and aperture in said rotor hub are configured to interference fit tolerances and then joined providing a joint having balance retention. An interference fit joint between a shaft extending from a bearing assembly and a rotor hub is also disclosed, wherein the joint is completed without using any mechanical fasteners or metallurgical bonding, yet can carry all of the operational loads placed on the joint.