Abstract: A sensor array having a lattice topology includes interconnects having an electrically-conductive layer sandwiched between two dielectric layers, the interconnects defining first-axis interconnects, second-axis interconnects, and interconnect junctions, sensor nodes located on associated interconnect junctions thereby defining an associated first-axis line and second-axis line, a sensor on an associated sensor node, a primary first-axis interconnect interface that is electrically connected to the first-axis interconnects, and a primary second-axis interconnect interface that is electrically connected to the second-axis interconnects. Each sensor node includes a first electrode that is electrically connected to an associated first-axis line, a second electrode that is electrically connected to an associated second-axis line, and a bypass bridge that electrically isolates the associated second-axis line from the associated first-axis line.

Abstract: A compressor case for a gas turbine engine includes an annular body that extends circumferentially around a center axis and extends axially along the center axis. A first bleed manifold is formed on an outer surface of the annular body and encloses a first plenum. A second bleed manifold is formed on the outer surface of the annular body and is axially aft of the first bleed manifold. The second bleed manifold encloses a second plenum. A bleed inlet extends through the annular body and into the first bleed manifold. Cooling passages are formed in the annular body, and each of the cooling passages extends from the first plenum to the second plenum and fluidically connects the first plenum to the second plenum.

Abstract: A rotor assembly of a gas-turbine engine may comprise a first rotor blade, a second rotor blade, a third rotor blade, a first platform sealing assembly and a second platform sealing assembly. The first platform sealing assembly may be disposed between a first platform of the first rotor blade and a second platform of the second rotor blade. The second platform sealing assembly may be disposed between the second platform and a third platform of the third rotor blade.

Abstract: A gas turbine engine includes a first spool, a second spool, a primary combustor, and a diffuser. The first spool includes a first compressor rotationally driven by a first turbine via a first shaft. The second spool includes a second compressor driven by a second turbine via a second shaft. The first compressor, the diffuser, and the primary combustor are arranged in series to provide a compressed airflow discharged from the first compressor to the primary combustor via the diffuser, which includes walls that diverge towards the primary combustor. The second compressor is fluidly coupled to the diffuser to receive at least a portion of the compressed airflow from the diffuser. The second turbine is fluidly coupled to the diffuser to discharge an expanded airflow to the diffuser.

Abstract: A rotor blade for a gas turbine engine including an inner ply layer group includes a flared region and an airfoil region, a neck region between the flared region and the airfoil region; and a platform shell that comprises a platform shell root region, a platform shell platform region, and a platform shell neck region between the platform shell root region and the platform shell platform region, the platform shell root region sheathes the flared region, and the platform shell neck region flares outwardly away from the neck region to intersect the platform shell platform region at a platform shell neck region perimeter around an airfoil opening perimeter.

Abstract: A furnace for removing a molybdenum-alloy refractory metal core through sublimation comprising a retort furnace having an interior; a sublimation fixture insertable within the interior of the retort furnace, the sublimation fixture configured to receive at least one turbine blade having the molybdenum-alloy refractory metal core; a flow passage thermally coupled to the retort furnace configured to heat a fluid flowing through the flow passage and deliver the fluid to the molybdenum-alloy refractory metal core causing sublimation of the molybdenum-alloy refractory metal core.

Abstract: A seal assembly includes a first feather seal with a first cooling hole extending through the first feather seal. The seal assembly also includes a second feather seal adjacent to the first feather seal. The second feather seal includes a second cooling hole extending through the second feather seal. The first cooling hole is positioned over at least a portion of the second cooling hole.

Abstract: A control system for limiting a power turbine torque of a gas turbine engine is disclosed. In various embodiments, the control system includes an engine control module configured to output an effector command signal to a gas generator of the gas turbine engine; a power turbine governor module configured to output to the engine control module a power turbine torque request signal; and a power turbine torque limiter module configured to output to the power turbine governor module a power turbine speed rate signal to limit a power turbine speed overshoot of the gas turbine engine.

Abstract: An after-fan system for an engine may comprise an after-fan turbine an electrical generator operationally coupled to the after-fan turbine, and an electric motor electrically coupled to the electrical generator. The electrical generator may be configured to generate an electrical current in response to rotation of the after-fan turbine. The electric motor may be configured to generate torque.

Abstract: A nacelle assembly of a gas turbine engine includes an annular structure defining a central axis, and having a radially inward surface and a radially outward surface, the radially inward surface at least partially defining a bypass duct. An aft portion of the radially inward surface at least partially defines an axially extending convergent-divergent exit nozzle. A secondary nozzle flap is radially spaced from the aft portion of the radially inward surface. The secondary nozzle flap and the aft portion of the radially inward surface define a secondary bypass duct therebetween. The secondary nozzle flap is operably connected to the annular structure such that the secondary nozzle flap is selectably movable relative to the aft portion of the radially inward surface, thereby changing a cross-sectional area of a secondary bypass duct exit.

Abstract: A gas turbine engine includes an engine spool with a turbine section, a propeller, a motor, and a variable drive arrangement. The variable drive arrangement mechanically couples the engine spool to the propeller, and includes a first gear connection rotationally coupled to the propeller, a second gear connection rotationally coupled to the engine spool, and a sprag clutch rotationally coupled to the first gear connection. The sprag clutch is configured to passively decouple the motor from the first gear connection when a torque input from the engine spool exceeds a torque input from the motor, and is configured to passively couple the motor to the first gear connection when the torque input from motor exceeds a torque input from the engine spool.

Abstract: A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis, the rotor disk comprising a slot within a rim of the disk and a spacer within the slot, the spacer having a clearance feature sized to permit rotation of a rotor blade from an insertion position to an installed position.

Abstract: A nickel-phosphorus alloy coating comprising a substrate having a surface; a nickel phosphorus alloy coating plated to the surface, the nickel phosphorus alloy consisting of phosphorus from 15.0 wt. percent to 20.9 wt. percent.

Abstract: A vane arm connection system for a gas turbine engine includes a vane stem having a head with flat contact surfaces; a vane arm having a claw, the claw having opposed arms having inwardly facing surfaces engaging the flat contact surfaces of the head; and a torque transfer member having a body defining an opening for engaging the flat contact surfaces of the head of the vane stem, and at least one arm extending from the body to contact the claw, whereby load from torque is transferred away from the inwardly facing surfaces.

Abstract: A surge control system includes a rotor system with at least one compressor section and at least one turbine section operably coupled to a shaft. The surge control system also includes sensors configured to collect sensor data from the rotor system, an electric motor operably coupled to the rotor system, and a controller. The controller is operable to detect surge event from the sensor data, determine an amount of power to apply to the rotor system, and increase the amount of power provided to the rotor system to recover from the surge event.

Abstract: A rotor disk assembly for a gas turbine engine includes a rotor disk that defines an axis; and a damper pivot support that extends from a rim of the rotor disk, the damper pivot support comprising an aperture that defines a pivot axis parallel to the axis. A method to dampen a rotor blade of a gas turbine engine, includes independently pivoting a first damper and a second damper of a damper assembly about the pivot axis to respectively contact a first blade and a second blade in response to centrifugal force below a platform of the respective first blade and second blade.

Abstract: A rotor blade for a gas turbine engine including an inner ply layer group that at least partially defines a base of a root region; a platform around the inner ply layer group that at least partially defines the base; and a platform over-wrap around the platform, the platform over-wrap at least partially defines the base.

Abstract: An engine control device may comprise a processor and a memory. The engine control device may be configured to modify a fuel flow based on a density of the fuel proximate a fuel nozzle. The engine control device may include a densimeter embedded in, or disposed proximate, the engine control device. The engine control device may include a temperature sensor embedded in, or disposed proximate, the engine control device. The engine control device may be electrically coupled to a fuel valve and/or configured to modulate the fuel valve based on a density of the fuel at the fuel valve.

Abstract: A cooling system for a plurality of conductive cables in a gas turbine engine includes a cooling source and an electric motor disposed in a tail cone. The cooling source may comprise an electric fan or an oil pump. The cooling source may be configured for active cooling of the plurality of conductive cables. The electric fan may be in fluid communication with ambient air during operation of the gas turbine engine.

Abstract: A hybrid electric gas turbine engine includes a fan section having a fan, a turbine section having a turbine drivably connected to the fan through a main shaft that extends along a central longitudinal axis, a gas generating core extending along a first axis that is radially offset from the central longitudinal axis, a first electric motor drivably connected to the main shaft, wherein the electric motor is colinear with the main shaft, and an electric compressor extending along a second axis that is radially offset from the central longitudinal axis, the electric compressor in fluid communication with the second turbine section.