Abstract: A turbine engine assembly is provided. The assembly includes a low-pressure turbine assembly including a first turbine section configured to rotate in a first rotational direction at a first rotational speed, and a second turbine section configured to rotate in a second rotational direction at a second rotational speed. The second rotational direction is opposite the first rotational direction and the second rotational speed is lower than the first rotational speed. The assembly also includes a first drive shaft coupled to the first turbine section, and a fan assembly including a first fan section coupled to the first drive shaft such that the first fan section rotates in the first rotational direction at the first rotational speed, and a second fan section coupled to the second turbine section such that the second fan section rotates in the second rotational direction at the second rotational speed.

Abstract: A turbomachine blade may include an airfoil and a shank coupled to the airfoil. The shank may include a cover plate having a first circumferential face and a second, opposing circumferential face. A radial cooling groove is positioned in the first circumferential face and is configured to allow a cooling fluid to pass from a first radial position to a second, different radial position relative to the platform. The radial cooling groove provides cover plate and shank cooling. In addition, the radial cooling groove may deliver fluid for purging gaps between blade platforms and cover plates, which prevents the ingestion of hot gas from the turbine flowpath.

Abstract: A gas turbine engine component includes a structure that provides a cooling passage. The structure has a turbulator with a pedestal joining opposing first and second surfaces. The turbulator includes first and second legs spaced apart from one another and adjoining the pedestal. The first leg adjoins the second surface, and the second leg adjoins the first surface.

Abstract: A radial position control assembly for a gas turbine engine stage includes a case structure. A supported structure is operatively supported by the case structure. A support ring operatively supports the supported structure. The supported structure and the support ring have different coefficients of thermal expansion. A sealing structure is adjacent to the supported structure. The support ring maintains the supported structure relative to the sealing structure at a clearance during thermal transients based upon a circumferential gap between adjacent supported structure and based upon a radial gap between the support ring and the supported structure. A pin supports the supported structure relative to the case structure and is configured to isolate the support ring from loads on the supported structure.

Abstract: A vane includes a vane body extending from a root to an opposed tip along a longitudinal axis and first and second baffle bodies. The vane body defines a leading edge and a trailing edge, and a cavity defined between the leading edge, the trailing edge, the root and the tip. The vane body includes at least one vane rib defined between the leading edge and the trailing edge inside the cavity. The first baffle body is defined in one of a leading edge portion and a trailing edge portion of the cavity. The second baffle body is defined in a middle portion of the cavity.

Abstract: The gas turbine component showerhead cooling hole layouts described herein include minimal lateral cooling hole exit diffusion on the middle showerhead cooling hole rows and interior facing sides of outer rows. In this way, rows of cooling holes may be placed close together. Stated another way, the outer showerhead cooling hole rows substantially only include lateral cooling hole exit diffusion in the direction away from the other rows to again allow the rows to be placed close together.

Abstract: A turbomachine blade comprising a blade tip and a metallic leading edge having a plurality of shear zones angled to the blade tip, wherein the shear strength of the shear zones is less than the shear strength of the remainder of the leading edge, such that in the event of an impact shear is initially initiated at the shear zones.

Abstract: An improved composite structure and method for making same has been provided. The provided improved composite structure has locally strengthened areas within a reinforcement region. The locally strengthened areas within the reinforcement region have load distribution devices to redistribute load in order to (i) locally strengthen an area around damage induced by an initial momentary and direct transmitted load, and (ii) limit growth and propagation of damage induced by an initial momentary and direct transmitted load during a subsequent unbalance load. The improved composite structure reduces the impact of the fan blade out phenomenon in a weight efficient manner.

Abstract: A balance piston seal assembly for a balance piston of a compressor is provided. The balance piston seal assembly may include a balance piston seal, a stationary support, and a gripping assembly disposed between the balance piston seal and the stationary support. The balance piston seal may be configured to be disposed about the balance piston such that an inner radial surface of the balance piston seal and an outer radial surface of the balance piston define a radial clearance therebetween. The stationary support may be configured to be coupled or integral with a casing of the compressor. The gripping assembly may be configured to secure the balance piston seal with the stationary support and to maintain concentricity between the balance piston seal and the balance piston during thermal radial expansion of the balance piston seal relative to the balance piston.

Abstract: A rotor for a gas turbine engine includes a cold shell, a hot shell, and a spoke. The spoke is connected to and extends radially outward from the cold shell. The hot shell is connected to the cold shell by the spoke and includes an axially extending outboard segment and an axially extending inboard segment. The outboard segment is connected to the inboard segment and the inboard segment is disposed radially inboard of the outboard segment for sealably engaging a stator blade shroud.

Abstract: Example implementations relate to a fan cover with a plurality of openings. For example, a device may include a fan cover to a fan device and a heat conduction device, a portion of the heat conduction device being coupled to an inside surface of the fan cover. The fan cover provides airflow for heat dissipation in a computing device and surrounds at least a portion of the fan device. The fan device has a first side associated with an axis about which an impeller of the fan device is capable of rotating and a second side associated with blade edges of the impeller of the fan device. The fan cover includes a flat portion to surround the first side and a curved portion to surround the second side, the curved portion having a plurality of openings such that airflow produced by the fan device is capable of passing through.

Abstract: A cooling system for gas turbine engines includes a turbine rotor compartment defining a cooling air plenum. A plurality of turbine discs are rotatably housed within the rotor compartment. A cooling air inlet is in fluid communication with the plenum. Each turbine disc includes a cooling outlet in fluid communication with the plenum for cooling the rotor compartment.

Abstract: A system for supplying pressurized air installed in an aircraft turbine engine, configured to supply pressurization air to a portion of the aircraft for using compressed air from pressurization air collected from a portion for collecting compressed air, characterized in that it comprises a collection port formed on a casing of the compressed air collection portion, a collection member coupled to the collection port, a through port of the collection member formed on a casing of a compartment of the turbine engine, said casing being subjected to small movements relative to the casing of the compressed air portion, the collection member crossing the through port with a freedom of movement relative to the latter during said small movements, a high-pressure space crossed by the collection member, located between the casing of the compressed air collection portion and the casing of the compartment.

Abstract: A stator of a low-pressure compressor of an axial-flow turbomachine. The stator includes an annular row of stator blades including radial extremities which pass through the openings of an internal shroud, and which include radial retaining slots having tapers formed by hooks. The stator includes a ring for retaining the blades on the internal shroud. The ring is curved circumferentially in order to be inserted into a plurality of retaining slots and exhibits the form of a strip having an arched transversal profile which is in abutment against the tapers, in such a way as to maintain the ring in the interior of the slots. The shroud includes an annular layer of abradable material made from silicone, which encloses the ring in such a way as to block the curvature of the arched profile of the ring in order to prevent it from disengaging from the tapers of the slots.

Abstract: A fitting according to an exemplary aspect of the present disclosure includes, among other things, a main body portion having a first axial end and a second axial end opposite the first axial end, an inlet provided at the first axial end, and an outlet adjacent the second axial end. The outlet includes a plurality of circumferentially spaced-apart slots formed through the main body portion, and each slot has an area within a range of 0.52 and 0.59 square inches.

Abstract: A hybrid compressor blade having an airfoil portion and a root portion that includes an outer shell comprised of one or multiple types of material that are each located at a predesignated section on the compressor blade, having at least one transition region between two different sections. The transition region is comprised of one or multiple layers with a compositional gradient based upon materials in the neighboring sections to provide a gradual transition from one section to another.

Abstract: An airfoil structure for a gas turbine engine includes an airfoil which includes a leading edge and a trailing edge. A platform is located adjacent a first end of the airfoil and includes a core passage that extends through the platform, a mate-face for engaging an adjacent airfoil structure and a set of impingement cooling holes in communication with the core passage that extend through the mate-face adjacent the trialing edge of the airfoil.

Abstract: A flow path component for a gas powered turbine includes a flow path component body having cross sectional profile having a leading edge and a trailing edge. The leading edge is connected to the trailing edge by a first side and by a second side opposite the first side. A serpentine cooling passage includes a plurality of segments, each of the segments being generally radially aligned. A first subset of the segments is disposed along one of the first side and the second side of the cross sectional profile, and a second subset of the segments spans the flow path component body from the first side to the second side. A cavity is positioned internal to the flow path component body. The cavity is at least partially shielded from one of the first side and the second side by at least one of the plurality of segments.

Abstract: A fan containment system for fitment around an array of radially extending fan blades mounted on a hub in an axial gas turbine engine. The fan containment system comprises an annular casing element for encircling an array of fan blades. An annular fan track liner is positioned substantially coaxial to the annular casing element. A hook projects in a generally radially inward direction from the annular casing element and is positioned axially forward of an array of fan blades when the fan containment system is fitted around said fan blades. Tabs protrude from the hook in a generally rearward direction and are spaced circumferentially along the hook. Fasteners connect the fan track liner to the tabs. The fasteners are configured so as to permit movement of the fan track liner towards the annular casing element when the fan track liner is impacted by a released fan blade.

Abstract: A rotor blade may generally include a shell forming an outer skin of the blade, with the shell defining a chordwise curvature. The rotor blade may also include a spar cap extending within the shell along a spanwise direction of the blade. The spar cap may be formed from an assembly of pre-cured laminate plates. In addition, the rotor blade may include an interior shelf positioned directly between the shell and the spar cap. The interior shelf may include an outer surface extending adjacent to the shell and an inner surface opposite the outer surface. The outer surface may define a curved profile generally corresponding to a portion of the chordwise curvature of the shell and the inner surface may define a planar surface along which the spar cap extends in a chordwise direction of the blade. The interior shelf may correspond to a pre-fabricated insert for the blade.