A fan casing arrangement for a gas turbine engine having a propulsive fan, the arrangement having a fan case and a fan track liner and being configured to circumscribe the fan, wherein the fan track liner is provided around the inside of fan case so as to adopt a radial position between the fan and the fan case. The fan track liner includes an elongate member which is helically-wound against the inside of the fan case in a plurality of turns. A method of installing a fan track liner in a fan casing arrangement for a gas turbine engine having a propulsive fan, the method involving: providing a fan case to circumscribe the fan, providing a flexible and elongate liner member, and helically winding the liner member against the inside of the fan case in a plurality of turns to define at least part of the fan track liner.

Oil slinger systems include a seal runner comprising an annular radial member having a radius (R) and an outer axially extending member having an axial length (L), wherein a proximal surface of the outer axially extending member comprising a plurality of helical grooves. Methods of radial convective cooling include pumping a cooling liquid through the oil slinger system and convectively cooling the oil slinger.

A core structure (10) includes a first core element (16) including a leading edge section (30), a tip section (32), and a turn section (34) joining the leading edge and tip sections (30, 32). The first core element (16) is adapted to be used to form a leading edge cooling circuit (102) in a gas turbine engine airfoil (100). The leading edge cooling circuit (102) includes a cooling fluid passage (104) having a leading edge portion (106) formed by the first core element leading edge section (30), a tip portion (108) formed by the first core element tip section (32), and a turn portion (110) formed by the first core element turn section (34). Each of the leading edge portion (106), the tip portion (108), and the turn portion (110) of the cooling fluid passage (104) are formed concurrently in the airfoil (100) by the first core element (16).

A seal system has: a first member; a seal carried by the first member and having a seal face; and a second member rotatable relative to the first member about an axis. The second member has: a seat on a first piece of the second member, the seat having a seat face in sliding sealing engagement with the seal face; and a radially outwardly closed collection channel for collecting centrifuged oil; a second piece encircling and attached to the first piece and having a circumferential array of apertures; and cooperating with the first piece to define a plenum; and a flowpath from the collection channel passing radially outward axially spaced from the seat face to cool the seat face and passing axially away from the seat face in the plenum.

A centrifugal compressor for compressing a fluid in a gas phase or a supercritical phase includes: a rotational shaft; an axial flow path extending along an axial direction of the centrifugal compressor; a radial flow path communicating with the axial flow path and extending along a radial direction of the centrifugal compressor on a downstream side of the axial flow path; an impeller at least partially disposed in the radial flow path and configured to rotate together with the rotational shaft to increase pressure of the fluid flowing in the radial flow path; and a pre-compression unit disposed in the axial flow path at a position distant from a leading edge of the impeller on an upstream side of the leading edge and configured to increase the pressure of the fluid in advance before the fluid is introduced to the leading edge.

Disclosed herein is a gas turbine disk that includes a cooling target, and a disk unit having a main passage that is open to supply cooling air to the cooling target, and a plurality of unit passages that are open at an end of the main passage while each having a predetermined size.

A helical cross flow pulse detonation engine.

A turbocharger contains a turbine housing having an accommodating region for a turbine rotor disk of the turbocharger, which accommodating region is arranged centrally with respect to a turbine housing axis, and a turbine spiral channel, which tapers helically toward the accommodating region. A wastegate valve, having a spindle arm and a valve plate arranged on the spindle arm, or a variable exhaust-gas guiding device, having bearing disks and guide vanes, is arranged in the turbine housing. At least one of the: turbine housing, spindle arm and valve plate, or bearing disks and guide vanes, has a steel material for high-temperature applications. The material composition of which contains, in addition to iron, Fe, at least the following alloying constituents in amounts within the specified limits in weight percent: carbon: 0.4-0.5%; silicon: 1.25-1.75%; manganese: 3.0-12.0%; chromium: 19.5-20.5%; nickel: 5.0-6.0%; niobium: 1.00-1.5%. The material composition ensures sufficient temperature resistance of the components.

A catalyst chamber with a catalyst bed embedded therein for a monopropellant thruster of a rocket engine. The catalyst chamber comprises an inlet having a first cross-sectional area through which a propellant can be introduced into the catalyst chamber and an outlet having a second cross-sectional area through which the propellant and/or resulting reaction products can be introduced into a combustion chamber of the thruster. The outlet is connected to the inlet via a catalyst volume of the catalyst chamber. At least one helical wall member is arranged within the catalyst chamber and is dividing the catalyst volume into two or more segments such that an effective length of the catalyst bed of each segment passed through by the propellant and/or its reaction products is larger than a geometrical length of the catalyst chamber defined between the inlet and the outlet along a direction of extension of the catalyst chamber.

A gas turbomachine includes a compressor portion, a turbine portion operatively connected to the compressor portion, and a combustor assembly fluidically connected to each of the compressor portion and the turbine portion. A turbine casing includes a body having an outer surface and an inner surface. A clearance control system includes a plurality of fluidically connected fluid channels extending through the turbine casing. The plurality of fluidically connected fluid channels includes a first fluid channel configured to direct a fluid flow in a first axial direction, a circumferential fluid channel configured to direct the fluid flow in a circumferential direction, and a second fluid channel configured to direct the fluid flow in a second axial direction substantially opposite the first axial direction. The first fluid channel includes a first outlet passing through the inner surface, and the second fluid channel including a second outlet passing through the inner surface.