Abstract: A pre-diffuser for a gas turbine engine includes an exit guide vane ring having a multiple of exit guide vanes defined around an engine longitudinal axis; a hot fairing structure adjacent to the exit guide vane ring to define a multiple of diffusion passages around the engine longitudinal axis; an outer radial interface between a radial outer surface of the hot fairing structure and the exit guide vane ring, the outer radial interface being a full hoop structure; and an anti-rotation feature between the hot fairing structure and the exit guide vane ring, the anti-rotation features inboard of the multiple of diffusion passages.

Abstract: A method and system for operating a gas turbine engine coupled to an aircraft propeller are described herein. The method comprises detecting a command for unfeathering the propeller, inhibiting shaft shear detection logic in response to detecting the command for unfeathering the propeller, detecting completion of the unfeathering of the propeller, and enabling the shaft shear detection logic in response to detecting the completion of the propeller unfeathering.

Abstract: A seal assembly includes a first and second seats. The first seat includes an axial portion extending circumferentially around a center axis and extending axially from a first end to a second end. A radial portion extends circumferentially around the axial portion and extends radially from the axial portion between the first end and the second end. A cavity is formed in the radial portion and extends circumferentially around the axial portion. A second seat includes a radial segment extending circumferentially around the center axis and extending radially from an inner surface to an outer surface of the radial segment. An axial segment extends circumferentially around the first end of the first seat, and extends axially from the radially segment to a distal end. The axial segment includes a tapered outer surface that tapers radially inward toward the distal end. The distal end is received in the cavity.

Abstract: An aerofoil component (10) defining an in use leading edge (16) and a trailing edge (18), the leading edge (16) comprising at least one serration (22) defining a sawtooth profile projecting in a generally upstream direction, each serration (22) comprising a plurality of rectangular-wave projections (26) extending in an upstream chordwise direction from the sawtooth profile.

Abstract: A turbine (1) for a conduit comprising a turbine unit (13) including a rotor (3), characterized in that the turbine (1) comprises a speed limiting device for said turbine unit (13), which device comprises at least one flexible element (4) connected to the turbine unit (13), so as to limit the flow rate of the fluid passing through the rotor (3) in order to limit the rotational speed of the rotor (3) of the turbine unit (13).

Abstract: The present invention relates to a guide vane ring element for a turbomachine, such as a gas turbine, with at least one drilled hole for the at least partial uptake of a flange of a rotatable guide vane, wherein an edge of the drilled hole on the vane element side is designed convexly along the edge in a first circumferential portion and, in particular, has a radius.

Abstract: A blade member includes a leading edge, a trailing edge, and a blade body extending therebetween. The blade member also includes a dovetail, a transition region, and a transition relief. The dovetail includes a dovetail chord line extending between an axially forward face of the dovetail and an axially aft face of the dovetail. The dovetail chord line has a dovetail chord length. The transition region is formed between the blade body and the dovetail. The blade body and the transition region form a blade body transition chord line at the interface of the blade body and the transition region which has a blade body transition chord length. The transition relief includes a geometric relief in at least one of a forward and an aft end of the transition region and the dovetail. The dovetail chord length is less than the blade body transition chord length.

Abstract: The invention relates to an intermediate housing hub (2) comprising a discharge stream conduit (18) having a given length (L1), and discharge fins (22) extending in the discharge stream conduit (18) and having a given chord length (L2), the chord length (L2) of at least one of the discharge fins (22) being at least equal to 50% of the length of conduit (L1).

Abstract: Aircraft turbine engine comprising a low-pressure spool that comprises a low-pressure shaft (24), means (44) for taking off power from said low-pressure shaft, and a fan (28) that is driven by said low-pressure shaft by means of a reduction gear (32), said reduction gear comprising at least one first element (50) that is connected to said low-pressure shaft for conjoint rotation, at least one second element (56) that is connected to said fan for conjoint rotation, and at least one third element (52) that is connected to a stator casing of the turbine engine, characterised in that said at least one third element is connected to said stator casing by disengageable connection means (60), and comprising at least one member that can move from a first position in which said at least one third element is fixedly connected to said stator casing into a second position in which said at least one third element is separated from said stator casing and is free to rotate about said longitudinal axis.

Abstract: An airfoil for a working fluid path of a turboengine extends along a spanwidth direction from a base to a tip. An aerodynamic body thereof includes a suction side surface, a pressure side surface, a leading edge, a trailing edge and a tip, the tip of the aerodynamic body having a tip cross section and a cross-sectional contour circumscribing the tip cross section. A rim extends to the tip of the airfoil and follows the cross-sectional contour on the pressure side, the suction side and extends over the leading edge of the airfoil, the rim delimiting a tip cavity which is open at the tip of the airfoil. The rim is further open at the trailing edge of the airfoil such that the tip cavity is open at the trailing edge of the airfoil.

Abstract: The gas turbine engine can have a first and second rotary components structurally joined to one another via a connector, a first spigot fit between the connector and the first rotary component, the first spigot fit forming an interference fit at a first operating condition and forming a gap at a second operating condition, a second spigot fit between the connector and the first rotary component, the second spigot fit forming an interference fit at the second operating condition and forming a gap at the first operating condition, the gas turbine engine being further configured to form a radial interference fit between the connector and the second rotary component in both the first and second operating conditions.

Abstract: A shroud assembly for gas turbine engine defining an axial direction, a radial direction, and a circumferential direction includes a shroud segment extending substantially along the circumferential direction and including a forward wall and an aft wall spaced along the axial direction from one another and defining a cavity therebetween; and a hanger assembly including a first attachment member and a seal member, the first attachment member attached to one of the forward wall or the aft wall of the shroud segment, and the seal member positioned within the cavity and interference fit against the other of the forward wall or the aft wall of the shroud segment.

Abstract: A turbine assembly adapted for use in a gas turbine engine is disclosed in this paper. The turbine assembly includes a turbine vane comprising ceramic matrix composite materials configured for use in high temperature environments. The turbine assembly further includes a vane-stage support for holding the turbine vane and other components in place relative to a turbine case.

Abstract: A vane ring for a gas turbine engine component includes a multiple of vanes that extend between an inner vane platform and an outer vane platform, each of the multiple of vanes contains an airfoil cooling circuit that receives cooling airflow through a feed passage located in the outer vane platform and an extension from the outer vane platform, the extension comprises a metering passage in communication with the feed passage.

Abstract: A cooling system includes a fan and a system component. The fan includes a plurality of fan blades and configured to rotate in a fan direction. The system component is located downstream of the fan, and includes a cutout for passing of airflow from the fan, and a bridge spanning the cutout. The bridge includes a center section and at least one arm section extending from the center section to an edge of the cutout along a curved path offset towards the fan direction.

Abstract: A waste gate valve for a turbocharger is provided that improves control responsivity, wear resistance and the flow uniformity of diverted exhaust gas by improving the operational structure thereof. The waste gate valve includes a turbine housing in which a portion of a bypass passage is formed along the inner circumferential surface of a discharge passage. A mixer ring includes bypass apertures therein that are arranged along the bypass passage formed in the inner circumferential surface of the discharge passage. Additionally, a control valve is configured to open or close the bypass apertures based on movement thereof.

Abstract: A case assembly for a turbine section has a structural outer case structurally connected to a gas generator upstream of the outer case and to an exhaust case downstream of the outer case. The assembly further comprises a containment ring mounted within the outer case and surrounding a plurality of axially spaced-apart turbine stages. An inner surface of the containment ring defines a plurality of shroud receiving portions. An annular plenum is defined between the outer case and the containment ring. The plenum has an inlet connected to a source of pressurized coolant. Outlets are circumferentially and axially distributed and defined through the containment ring. The outlets provides flow communication between the annular plenum and the plurality of axially spaced-apart turbine stages.

Abstract: A connection and fastening unit for components of a lightning protection system, the connection and fastening unit being realized for the purpose of being integrated in a wind turbine rotor blade and for being connected to the components of the lightning protection system, wherein the connection and fastening unit includes a cavity which can be closed in an airtight manner and in which a connector for connection to at least one of the components of the lightning protection system is arranged.

Abstract: An elastic flapping hinge for connecting a rotor blade to a rotor hub of a rotary wing aircraft, comprising an elastic flapping hinge member arrangement that includes a hub attachment area for attachment to the rotor hub, a connection area for attachment to the rotor blade, and an elastic flapping hinge area that is arranged between the hub attachment area and the connection area and adapted to allow flapping movements, the elastic flapping hinge member arrangement comprising at least two elastic flapping hinge members having a first bending stiffness for flapping movements and a second bending stiffness for lead-lag movements, the first bending stiffness being smaller than the second bending stiffness, wherein the at least two elastic flapping hinge members diverge from each other in the elastic flapping hinge area by a predetermined divergence angle.

Abstract: The present invention relates to a heat dissipation fan, where the heat dissipation fan includes a fan frame and a fan blade, where the fan blade includes a hub and a plurality of double-layer blades evenly disposed around the hub, and a primary air intake area is formed between the double-layer blades; the double-layer blade includes a primary blade and a secondary blade, where the secondary blade is fixed on the primary blade, and a secondary air intake area is formed between the primary blade and the secondary blade; and when the heat dissipation fan works, an airflow enters the primary air intake area and the secondary air intake area of the heat dissipation fan through an air inlet of the heat dissipation fan, and forms a first airflow in the primary air intake area and a second airflow in the secondary air intake area, and the first airflow and the second airflow flow out through an air outlet of the heat dissipation fan.