Abstract: The invention relates to an assembly for a turbomachine, comprising a stator (1) and a rotor (2) rotatable relative to the stator (1) about an axis, the rotor (2) comprising blades each ×comprising a vane (3) connected to a radially inner platform (5), a block of abradable material (6) extending radially inwardly from the radially inner platform (5) the stator (1) having a shroud comprising an annular area (9), at least one lug (12) extending radially outwardly from said annular area (9), the radially outer end of the lug (12) cooperating with the block of abradable material (6).

Abstract: The disclosure relates to a centrifugal compressor and a diffuser; the diffuser includes a pressure drive mechanism, a first diffuser part, a second diffuser part and a movable diffuser part movable diffuser part being connected with the pressure drive mechanism and movably arranged on one of the first diffuser part and the second diffuser part; moreover, the movable diffuser part gets close to or leaves away from the other one of the first diffuser part and the second diffuser parts under the action of a pressure medium in the pressure drive mechanism so as to adjust a width of a pressure diffusion flow channel and prevent adverse phenomena such as gas flow stall and surging.

Abstract: The present invention provides an airfoil 110 with the squealer tip cooling system 50 for a turbine blade 100 at the blade tip 113, wherein the squealer tip cooling system 50 comprises a cooling passage 170 arranged within a squealer tip 117, wherein the cooling passage 170 at least partly extends toward a terminal end 74 of the squealer tip 117, and a pocket 172 at a lateral surface 75, 76 of the squealer tip 117, open externally and extending inwardly at least partly across the cooling passage 170. The pocket 172 intersects the cooling passage 170 and the pocket 172 comprises an impingement surface 70 facing the cooling passage 170, on which a cooling medium expelled through the cooling passage 170 impinges before being discharged externally through the pocket 172.

Abstract: A seal assembly for a gas turbine engine having a seal formed of a carbon material; and a seal seat positioned for rotation relative to the seal, wherein the seal and the seal seat each have a sealing surface which together define a sliding seal, and further having a carbon film on the sealing surface of the seal seat.

Abstract: A turbine stage includes an array of airfoils spaced apart circumferentially to define a flow passage therebetween for channeling a working medium. The airfoils extend radially outward from an inner endwall located at a hub side thereof. The inner endwall is inclined at an angle to an engine axis such that the flow passage is divergent from an upstream side to a downstream side. The inner endwall is non-axisymmetric about the engine axis, having a mid-passage bulge located between circumferentially adjacent first and second airfoils. The bulge has a peak at a position between 20-60% CaxID and at a position between 30-70% pitchID.

Abstract: A turbine engine blade having an aerodynamic surface extending a first direction between a leading edge and a trailing edge and in a second direction, perpendicular to the first direction, between a root of the blade and a tip of the blade, the aerodynamic surface being made of a fiber-reinforced organic matrix composite material, and a metallic structural reinforcement bonded by an adhesive joint to the leading edge whose shape it follows and which has over its entire height a substantially V-shaped section with a base extended by two lateral flanks having a thinned profile at free ends directed toward the trailing edge, the adhesive joint being locally supplemented below the free ends of the lateral flanks by an elastomeric polymer introduced in the form of solid particles into the adhesive joint and adhered to the aerodynamic surface and/or the free ends of the lateral flanks during a polymerization phase.

Abstract: A labyrinth seal, has: rotating and static components rotatable relative to one another relative to a central axis, the rotating component securable to a shaft via a tight fit engagement at an engagement location on the rotating component, the static component securable to a housing; teeth protruding from one of the rotating and static components towards a seal land defined by the other one of the rotating and static components; and clearances between the teeth and the seal land, a first clearance of the clearances greater than a second clearance of the clearances, the first clearance located closer to the engagement location of the rotating component than the second clearance.

Abstract: A turbofan gas turbine engine includes heat exchanger module, fan assembly, compressor, turbine and exhaust modules. The fan includes a plurality of fan blades. The heat exchanger in fluid communicates with the fan assembly by an inlet duct, and the heat exchanger includes a plurality of radially-extending hollow vanes arranged in a circumferential array, with a channel extending axially between each pair of adjacent hollow vanes. An airflow entering the heat exchanger is divided between a set of vane airflows and a set of channel airflows. Each vane airflow has a vane mass flow rate FlowVane, and each channel air flow has a channel mass flow rate FlowChan. Each hollow vane includes, an inlet, heat transfer, and exhaust portions, with the inlet portion comprising a diffuser element and the heat transfer portion including at least one heat transfer element. The diffuser element causes FlowVane to be lower than FlowChan.

Abstract: Hybrid material components and methods of forming hybrid material components are provided. For example, a hybrid material component comprises a plurality of metallic tows and a plurality of non-metallic tows. Each metallic tow of the plurality of metallic tows is surrounded by a portion of the plurality of non-metallic tows such that the plurality of metallic tows are embedded within the plurality of non-metallic tows. An exemplary hybrid material component is a containment assembly of a gas turbine engine. Methods of forming such components include forming a hybrid material; laying up a plurality of layers of the material to form a layup; and processing the layup such that a plurality of metallic tows and a plurality of non-metallic tows are co-cured.

Abstract: A gas turbine engine guide vane heat exchanger has guide vane heat exchanger including electroformed fluid channels in electroformed heat exchanger tubes or a heat exchanger core disposed within airfoil. Non-flammable heat conducting liquid or non-metallic foam may fill space between tubes or core and airfoil. Fluid circuit may include channels within electroformed heat exchanger tubes or the heat exchanger core and extend from inlet manifold to outlet manifold for directing fluid or oil through channels and include fluid or oil supply inlet connected to inlet manifold for receiving the fluid or oil flowed into inlet manifold and a fluid or oil supply outlet connected to fluid or oil supply outlet for discharging fluid or oil flowed out of fluid or oil outlet manifold. Heat exchanger tubes or heat exchanger core, inlet manifold, outlet manifold, supply inlet and supply outlet may be integrally and monolithically electroformed together.

Abstract: A rotor and a turbo machine including the same are provided. The rotor includes a disk including a disk slot, a blade including a root member inserted into the disk slot and an airfoil disposed radially outside the root member, and configured to form a cooling cavity between an inner surface of the disk slot and the root member, a lifting part installed in the cooling cavity and configured to press the root member in a radially outward direction, and a supporting part installed in the cooling cavity, disposed in contact with the lifting part, and configured to support the lifting part.

Abstract: A modular gearbox assembly for a wind turbine having improved up-tower serviceability includes a low-speed gear stage module, a separate, intermediate-speed gear stage module adjacent to the low-speed gear stage module, and a separate high-speed gear stage module adjacent to the intermediate-speed gear stage module. The gearbox assembly also includes a first flange removably connecting the intermediate and high-speed gear stage modules and a second flange removably connecting the intermediate and low-speed gear stage modules. Thus, the low-speed gear stage module converts a low-speed, high torque input from a rotor shaft of the wind turbine to a high-speed, low torque output for a generator of the wind turbine via the intermediate and high-speed gear stage modules. In addition, the first and second flanges allow for easy disassembly of the gear stage modules such that the various stages can be easily repaired, replaced, and/or inspected.

Abstract: An impingement baffle for directing a cooling fluid onto a target surface includes a baffle body having a first end opposite a second end, and a first side opposite a second side. The second side is spaced a distance apart from the target surface, with the distance varying from the first end to the second end. The baffle body defines impingement holes that extend through the baffle body from the first side to the second side. The impingement holes are spaced apart along the baffle body to receive the cooling fluid. The impingement baffle includes tubular extensions coupled to the second side. Each tubular extension is in fluid communication with a respective one of the impingement holes to direct the cooling fluid onto the target surface. Each tubular extension extends for a length from the second side, and the length of each tubular extension is based on the distance.

Abstract: A system and a method of designing a nacelle for a gas turbine engine having a longitudinal centre line. The nacelle includes an air intake, an internal surface, an azimuthal angle, and a plurality of intake lines. The air intake comprises, in flow series, an intake lip, a throat and a diffuser. The internal surface at least partially defines the air intake. The azimuthal angle is defined about the longitudinal centre line. The intake lines extend along the internal surface of the nacelle at respective values of the azimuthal angle. Each intake line axially defines the air intake along the longitudinal centre line at the respective value of the azimuthal angle. The internal surface of the nacelle between the plurality of intake lines at a given axial location along the longitudinal centre line is analytically defined by an equation.

Abstract: A gas turbine engine defining a longitudinal direction and a radial direction is provided.

Abstract: Methods and apparatus for real-time clearance assessment using a pressure measurement are disclosed. An example method includes determining a first and a second static pressure measurement at a first measurement location and a second measurement location, respectively, relative to the blade tip clearance, determining a normalized pressure measurement using the first and second static pressure measurements, generating a conversion curve to correlate the normalized pressure measurement with a clearance measurement, and adjusting active clearance control of the blade tip clearance based on a comparison of real-time in-flight pressure measurements to the conversion curve.

Abstract: A support assembly for a gas turbine engine is provided. The support assembly includes an outer casing, an inner structure, a strut and a mechanical fastener. The outer casing defines a longitudinal central axis. The inner structure includes a platform and a support portion disposed around the platform. The platform defines an aperture therethrough and the support portion defines a slot therethrough adjacent to the platform. The slot at least partially surrounds the platform and extends axially towards the aperture by an axial slot length along the longitudinal central axis. The strut extends generally radially from the inner structure to the outer casing. The mechanical fastener is received through the aperture of the platform.

Abstract: A rotating seal rotatable about a rotational axis is provided. The rotating seal includes a body having a first surface disposable to face the rotational axis and a second surface disposable to contact with a stationary element. The body defines a cooling channel including one or more entrance channels respectively extending from the first surface, one or more exit channels and a plenum. The plenum extends circumferentially through the body and has a hot side adjacent to the second surface. The plenum is fluidly interposed between the one or more entrance channels and the one or more exit channels whereby fluid exiting the one or more entrance channels and entering the plenum impinges against the hot side.

Abstract: A composite aerofoil may include an aerofoil body defining a leading edge and a trailing edge, wherein the body comprises a composite material including a plurality of relatively higher-modulus reinforcement elements, a plurality of relatively tougher polymer-based reinforcement elements, and a matrix material substantially encapsulating the plurality of relatively higher-modulus reinforcement elements and the plurality of relatively tougher polymer-based reinforcement elements. The plurality of relatively higher-modulus reinforcement elements are different from the plurality of relatively tougher polymer-based reinforcement elements. The disclosure also describes techniques for forming composite aerofoils.

Abstract: The present invention relates to a method for designing a flow channel for a turbomachine, in particular a gas turbine that comprises a guide vane cascade having a plurality of guide vanes, which are distributed in the peripheral direction, and flow passages, each of which is bounded by two successive guide vanes, and a support rib arrangement having at least one support rib, wherein a design of one of the flow passages is adapted to this support rib, that it is situated downstream of, in order to reduce a pressure loss and/or a vibrational stimulation.