Casings and methods for manufacturing casings are provided. For example, a casing defining radial, axial, and circumferential directions is provided. The casing comprises an annular inner wall and an annular outer wall, each extending along the axial direction, with the outer wall radially spaced apart from the inner wall. The casing also comprises an auxetic structure extending from the inner wall to the outer wall and including a plurality of lattice elements. Each lattice element extends circumferentially and radially from the inner to the outer wall, and the lattice elements are axially spaced apart from one another. The auxetic structure may define at least one aperture for fluid flow from one portion to another of the auxetic structure and/or may be configured to vary the thermal characteristics of the casing along the axial direction. The casing may be integrally formed as a single monolithic component, e.g., by additive manufacturing.

A bearing compartment seal for a gas turbine engine includes at least one seal ring defining an axis and having a radially inward facing sealing surface and a seal runner having a support constructed of a first material and an interface portion constructed of a second material. The interface portion includes a radially outward facing surface. A first coefficient of thermal expansion of the second material is at most approximately equal to a second coefficient of thermal expansion of the at least one seal ring.

A system and method for rotor bow mitigation for a gas turbine engine are provided. An elapsed time since a shutdown of the engine and an idle operation time of the engine prior to the shutdown are determined. A rotor bow mitigation period is determined based on the elapsed time and the idle operation time and, prior to initiating a start sequence of the engine, the engine is motored for a duration of the rotor bow mitigation period.

A mechanical connection system is provided that includes a ceramic skin panel, at least one bushing, and at least one bolt. The ceramic skin panel includes a first surface, a second surface, and at least one aperture extending between the first surface and the second surface. The aperture has an axial centerline, a chamfered portion contiguous with the first surface, and an axial portion contiguous with the second surface. The chamfered portion includes a tapered surface disposed at a first angle such that a line extending along the tapered surface intersects the axial centerline at a plane of the second surface. The bushing has a housing and a collar. The housing is configured to mate with the chamfered portion of the aperture disposed within the ceramic skin panel. The collar is configured for engagement with the housing.

An assembly of at least two members. One of the members supports the other member, the assembly defining an axial direction, a radial direction, and a circumferential direction, an inner member of the at least two members being received radially inside an outer member of the at least two members, wherein the inner member and the outer member am attached to each other by a support arrangement, the support arrangement including at least one floating support assembly as a displaceable coupling between an inner member support point provided at the inner member and an outer member support point provided at the outer member. A displacement of support points in a radial direction results in an interrelated relative displacement of the support points in an axial direction end vice versa.

An assembly for a turbojet, wherein the assembly includes an outer shroud and an inner shroud that are concentric, wherein the inner shroud is segmented and includes circumferential clearances between the segments thereof. The assembly additionally includes an annular row of stator vanes connecting the inner shroud to the outer shroud, a drive with a reduction ratio that is intended to be coupled to a fan, and a circuit for cooling and for lubricating the drive. The circuit is configured to heat up at least the outer shroud during the operation of the turbine engine such as to circumferentially reduce the circumferential clearances between the segments.

A coating includes: at least 34.9 percent by mass silicon dioxide; at least 9.1 percent by mass aluminum oxide; and at least 16.1 percent by mass yttrium oxide.

There is provided a manufacturing method of a turbine casing capable of easily realizing improvement of reliability. A manufacturing method of a turbine casing according to an embodiment is a manufacturing method of a turbine casing which includes an outer casing formed of ferritic heat resistant steel and an inner casing disposed inside the outer casing and formed of austenitic heat resistant steel, and in which an exhaust hood to which a working medium after performing work in turbine stages is exhausted, is covered by the inner casing. Here, the inner casing is manufactured by using members produced by at least either forging or rolling.

A cartridge for a fan case of a gas turbine engine includes an inlet acoustic liner section integrated with a thermally conforming liner section.

A passive clearance control limits thermal expansion between stator components relative to rotor components. A control ring controls clearance in a passive manner and is located on or adjacent to stationary components which thermally expand during engine operation. The control ring is formed of material having low coefficient of thermal expansion such as CMCs (Ceramic Matrix Composites) and therefore limits, inhibits or restrains expansion of the adjacent stator components as temperatures increase. Limiting expansion of the stator component reduces rotor/stator clearances and limits parasitic leakage of fluid along the flowpath through the engine core.