The present disclosure relates to turbine blades adapted for use in gas turbine engines. In particular, this disclosure is directed to turbine blades that include components made from ceramic matrix composite materials and that incorporate abradable materials.

A fairing for a power generation machine, the fairing comprising: a first layer comprising a metallic material and defining a first surface and a second opposite surface, the first layer being configured to enable the fairing to be coupled to the power generation machine; and a second layer comprising a composite material and defining a first surface and a second opposite surface, the second layer being coupled to the first layer, the second surface of the second layer defining an external surface of the fairing.

An aircraft engine comprises a casing extending circumferentially around a central axis. A support flange is secured to the casing and extends circumferentially around the axis. The support flange has an inner and an outer flange face. A component is drivingly engaged by a shaft of the engine. The component has a connecting section extending around the central axis. The connecting section has a connecting face. The inner flange face is in abutment against the connecting face. A retaining ring extends circumferentially around the central axis. The retaining ring is in abutment against the outer flange face of the support flange. The retaining ring has a coefficient of thermal expansion, which is less than that of the support flange.

Platforms for airfoils in gas turbine engines are provided. The platforms include an aero-structure defining a gaspath surface and a non-gaspath surface, a first securing element arranged to receive a locking pin and secure the aero-structure to a rotor disk, and a first securing element structure extending between the non-gaspath surface and the first securing element, wherein the first securing element structure comprises at least two legs defining a hollow space between the two legs and the non-gaspath surface.

A seal system for a gas turbine engine includes a seal runner manufactured of a Molybdenum alloy material that provides a first coefficient of thermal expansion and a seal ring manufactured of a graphitic material that provides a second coefficient of thermal expansion greater than the first coefficient of thermal expansion.

An aircraft engine comprises a casing extending circumferentially around a central axis. A support flange is secured to the casing and extends circumferentially around the axis. The support flange has an inner and an outer flange face. A component is drivingly engaged by a shaft of the engine. The component has a connecting section extending around the central axis. The connecting section has a connecting face. The inner flange face is in abutment against the connecting face. A retaining ring extends circumferentially around the central axis. The retaining ring is in abutment against the outer flange face of the support flange. The retaining ring has a coefficient of thermal expansion, which is less than that of the support flange.

A method of controlling macroscopic strain of a porous structure includes contacting a porous structure with a modifying agent which chemically adsorbs to a surface of the porous structure and modifies an existing surface stress of the porous structure. Additional methods and systems are also presented.

A turboramjet has a housing with an intake and an exhaust. The housing houses a heat exchanger, a turbojet section and a ramjet section downstream of the turbojet section. The heat exchanger has an air path and a coolant path. The air path is configured to receive air from the air intake. The heat exchanger has a first section made from a first material and a second section made from a second material, the second material having a lower melting point and a lower density relative to the first material. A bypass air passage selectively bypasses the turbojet section to supply air to the ramjet section, and the coolant path uses fuel as a coolant and is configured to supply the fuel to the turbojet section.

A method of repairing a tip of an airfoil can comprise: removing a coating from the tip of the airfoil; preparing a repair area for bonding, the repair area including the tip and a portion of a concave side of the airfoil; applying an epoxy film and an abrasive component to the repair area; and bonding the epoxy film and the abrasive component to the repair area.

An engine bi-material joint includes a first flange composed of a first material and defining a first coefficient of thermal expansion, and a second flange composed of a second material and defining a second coefficient of thermal expansion. The second flange is different from the first material. An interface flange is engaged with the first flange and with the second flange. The interface flange defines a third coefficient of thermal expansion being equal to or less than the first coefficient of thermal expansion of the first flange. The third coefficient of thermal expansion is less than the second coefficient of thermal expansion of the second flange. The first coefficient of thermal expansion of the first flange is less than the second coefficient of thermal expansion of the second flange.