Abstract: A gas turbine engine includes a compressor, a cooling source, and a thermoelectric intercooler adapted for selective operation in response to operational states of the gas turbine engine.

Abstract: A gas turbine engine, a blisk and methods of manufacturing are disclosed. The blisk includes a disk and a plurality of blades. Each blade includes a platform, an airfoil, and a shank portion radially extending between the platform and the disk. The airfoil includes cooling features. Cooling cavities are defined between shank portions of adjacent blades and are in communication to the internal of blades via cooling feed channels. A coverplate is coupled to a downstream side of the blisk to cover the cooling cavities. A plurality of inner tabs extends from an inner edge of the coverplate to securely couple the coverplate to the disk. The disk and the plurality of blades may be integrally formed or casted as a single unit. The platform, the airfoil, and the shank portion may be formed integrally as a single unit and bonded to the disk to form the blisk.

Abstract: A gas turbine engine and a blisk are disclosed, including methods of making the same. The blisk may include a disk, spars of airfoils, platforms, and shanks of airfoils integrally formed as a unit. Such blisk features may be integrally formed by a casting process or additive manufacturing. The blisk includes cooled airfoils, such that the spar is hollow, and a cover panel is bonded to an outer surface of standoffs to define passages between the cover panel and the spar. Cooling air is fed to the hollow spar through a cooling feed channel defined in the disk or shank portion.

Abstract: A radial turbine rotor for a gas turbine engine may be provided. The radial turbine rotor may include a hub and a plurality of radial turbine blades. The hub may include an outer surface. The outer surface of the hub may include a plurality of discrete bonding surfaces. The radial turbine blades may be bonded to a corresponding one of the discrete bonding surfaces of the hub. Each of the radial turbine blades may be separate and distinct from the other radial turbine blades that are bonded to the radial turbine rotor. The corresponding discrete bonding surface may include a planer surface to receive the radial turbine blade.

Abstract: A compressor shroud assembly in a turbine engine having a dynamically moveable shroud for encasing a rotor segment. The rotor segment comprises a bladed disc. The compressor shroud assembly maintains a clearance gap between the shroud and the blade tips of the bladed disc. The assembly comprises a static compressor casing, a shroud mounted to the casing, and an actuator mounted to the casing. The shroud comprises a flowpath boundary member spaced radially inward from the casing, the flowpath boundary member being moveable relative to the casing in a radial direction. The actuator is coupled to the shroud for effecting the movement of the flowpath boundary member.

Abstract: An expandable exhaust cone assembly is described which is able to move from a collapsed position to an expanded position.

Abstract: A gas turbine engine includes a lubrication system, fuel system and thermoelectric heat exchanger adapted for selective operation in response to operational states of the gas turbine engine.

Abstract: A compressor for use in a gas turbine engine is disclosed herein. The compressor includes a case, vanes, and a vane actuation system. The vanes are arranged circumferentially adjacent to one another about a central axis inside the case and each one of the vanes is configured for rotation about a vane axis generally perpendicular to the central axis. The vane actuation system is configured to cause rotation of at least one of the vanes about the vane axis.

Abstract: An axial compressor comprises a plurality of compressor stages positioned axially adjacent one another within a casing. Each of the plurality of compressor stages comprise a rotor segment and a banded stator segment positioned axially adjacent the rotor segment. An annulus is formed between the casing and an outer flowpath ring of the stator segment. A vane of each stator segment comprises a member extending into the annulus.

Abstract: An axial flow compressor comprises a tubular casing which encases a rotatable shaft, a pair of rotor segments coupled to the rotatable shaft and each comprising a bladed disc, and a banded stator segment disposed between the pair of rotor segments and comprising a plurality of stator vanes extending between an outer flowpath ring and an inner flowpath ring. A method of assembling an axial flow compressor comprises installing a rotor segment inside a tubular compressor casing, installing a vane segment adjacent the installed rotor segment, and repeating the steps of installing a rotor segment and vane segment until a desired number of rotor segment and vane segment pairs are installed.

Abstract: An axial compressor comprises a plurality of compressor stages positioned axially adjacent each other within a casing. Each of the plurality of compressor stages includes a rotor segment and a banded stator segment. An annulus is formed between the casing and an outer flowpath ring of the banded stator segment. A pathway may be provided that establishes an air flowpath between the annulus and another annulus formed by an adjacent stage.

Abstract: An axial compressor comprises a plurality of compressor stages positioned axially adjacent within a casing. Each of the plurality of compressor stages comprise a rotor segment and a banded stator segment positioned axially adjacent the rotor segment. An annulus is formed between the casing and an outer flowpath ring of the stator segment. A pathway is provided that establishes an air flowpath between an entry pathway in a first stage and an exit pathway in a second stage.

Abstract: One embodiment of the present disclosure is a unique active seal system. The active seal system includes a rotor and a stationary seal component disposed adjacent to the rotor. The rotor has a rotating seal component and a first electrical generator element. The stationary seal component has a second electrical generator element and a piezoelectric portion in electrical communication with the second electrical generator element.

Abstract: A ceramic matrix composite (CMC) seal segment for use in a segmented turbine shroud for radially encasing a turbine in a gas turbine engine. The CMC seal segment comprises an arcuate flange having a surface facing the turbine and a portion defining a bore for receiving an elongated pin, with the bore having a length that is at least 70% of the length of the elongated pin received therein. The CMC seal segment is carried by the carrier by at least one of the elongated pins being received within the bore. The CMC seal segment portion defining a pin-receiving bore is radially spaced from the arcuate flange by a spacing flange extending radially outward from the arcuate flange.

Abstract: A blade for a gas turbine engine includes a fixed length member and a floating blade seal that is movable relative to the floating blade seal to change the length of the blade and vary the gap between the blade and an engine housing component.

Abstract: A gas turbine engine includes a cooling air system, a coolant system, and a thermoelectric heat exchanger adapted for selective operation in response to operational states of the gas turbine engine.

Abstract: A high-Mach engine includes a gas turbine core, an inlet assembly, and a transmission. The inlet assembly including an inlet turbine and the transmission configured to couple the inlet turbine and a core turbine of the gas turbine core cooperate to control air moving through the high-Mach engine.

Abstract: An inlet assembly for a high-mach engine includes a gas turbine core, an inlet turbine, and a core-flow director. When in a closed position, the core-flow director forces air to interact with the inlet turbine before entering the gas turbine core.

Abstract: An airfoil may include a spar comprising a passageway inside of the spar for a cooling fluid, a pedestal on an outer surface of the spar, and an inlet configured to direct the cooling fluid from the passageway to the outer surface of the spar. The airfoil may further include a coversheet, wherein an inner surface of the coversheet is positioned on the pedestal of the spar and an edge of the coversheet is positioned along an end of the spar. The inner surface of the coversheet, the pedestal, and the outer surface of the spar may define a cooling path from the inlet to an outlet at the edge of the coversheet. The outlet at the edge of the coversheet may be configured to direct cooling fluid onto the end of the spar, onto a footing, and/or onto a fillet positioned along an intersection of the footing and the spar.

Abstract: The present disclosure provides, in one embodiment, a unique gas turbine engine. Other embodiments include unique gas turbine engines, and unique active balancing systems; as well as apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and active balancing systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.