A seal assembly includes at least one blade outer air seal (BOAS) including a base portion that extends between a leading edge and a trailing edge. A forward wall and an aft wall extend radially outward from the base portion to a radially outer portion. The radially outer portion is spaced from the base portion and at least partially defines a passage with the forward wall, aft wall, and base portion. At least one attachment body is located at least partially within the passage. The at least one attachment body includes a leading edge connected to a trialing edge by a radially inner surface and a radially outer surface. The radially outer surface includes at least one blade outer air seal attachment surface for engaging the at least one BOAS. At least one post extends from the radially outer surface and has a blade outer air seal guide surface.

An assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a gas turbine engine component that has a first interface portion, and a support that has a mounting portion and a second interface portion, the mounting portion attachable to an engine static structure, a first retention feature that releasably secures the first interface portion to the support in a first installed position of the gas turbine engine component, and a second retention feature dimensioned to secure the first interface portion to the second interface portion in a second installed position of the gas turbine engine component. The first installed position differs from the second installed position, and one of first and second retention features is dimensioned to carry the gas turbine engine component in response to release of another one of the first and second retention features. A method of sealing for a gas turbine engine is also disclosed.

A method for coating a substrate includes spraying a combination of powders. The combination of powders includes: Hf.sub.0.5Si.sub.0.5O.sub.2; Zr.sub.0.5Si.sub.0.5O.sub.2; and, optionally, at least one of HfO.sub.2 and ZrO.sub.2. A molar ratio of said Hf.sub.0.5Si.sub.0.5O.sub.2 and HfO.sub.2 combined to said Zr.sub.0.5Si.sub.0.5O.sub.2 and ZrO.sub.2 combined is from 2:1 to 4:1. A molar ratio of said Hf.sub.0.5Si.sub.0.5O.sub.2 to said HfO.sub.2 is at least 1:3.

An apparatus is disclosed including a first and second article, a first interface volume disposed between and enclosed by the first article and second article, a cooling fluid supply, and at least one cooling fluid channel in fluid communication with the cooling fluid supply and the first interface volume. The first article includes a first material composition. The second article includes a second material composition. The at least one cooling fluid channel includes a heat exchange portion disposed in at least one of the first and second article downstream of the cooling fluid supply and upstream of the first interface volume. A turbine shroud is disclosed wherein the first and second articles are an outer and inner shroud. A turbine nozzle is disclosed wherein the first and second articles are an endwall and fairing.

A blade outer airseal has a body comprising: an inner diameter (ID) surface; an outer diameter (OD) surface; a leading end; and a trailing end. The airseal body has a metallic substrate and a coating system atop the substrate along at least a portion of the inner diameter surface. At least over a first area of the inner diameter surface, the coating system comprises an abradable layer and a thermal barrier layer between the abradable layer and the substrate; and the thermal barrier layer comprises a ceramic and metallic phases within the ceramic.

A turbine airfoil is provided with at least one insert positioned in a cavity in an airfoil interior. The insert extends along a span-wise extent of the turbine airfoil and includes first and second opposite faces. A first near-wall cooling channel is defined between the first face and a pressure sidewall of an airfoil outer wall. A second near-wall cooling channel is defined between the second face and a suction sidewall of the airfoil outer wall. The insert is configured to occupy an inactive volume in the airfoil interior so as to displace a coolant flow in the cavity toward the first and second near-wall cooling channels. A locating feature engages the insert with the outer wall for supporting the insert in position. The locating feature is configured to control flow of the coolant through the first or second near-wall cooling channel.

Disclosed is a blade for a turbomachine, comprising an outer shell and an inner core which is at least partially enclosed by the outer shell and has a higher porosity than the outer shell. The outer shell is formed by a ceramic body or a body made of a ceramic matrix composite material, and the inner core is formed by a fiber-reinforced ceramic or a fiber-reinforced ceramic matrix composite material.

A turbine intermediate casing for a gas turbine comprising outer and inner wall elements arranged in series in circumferential direction and delimiting a flow passage for exhaust gas; strut fairing elements arranged in radial direction between respective outer and inner wall elements; an outer casing encompassing the outer and inner wall elements; and an annular mounting structure encompassing a bearing region of a shaft of the gas turbine. The strut fairing elements comprise a space for accommodating carrier elements which extend from the annular mounting structure through the flow passage to the outer casing exhaust gas flowing through the flow passage being directed by the strut fairing elements around the carrier elements. The outer and inner wall and strut fairing elements comprise ceramic fiber composite materials, the outer and inner wall elements being connected to the outer casing and the annular mounting structure respectively.

A die cast system in which an external shell and an internal core usable to form a component of a gas turbine engine are formed together is disclosed. In at least one embodiment, the external shell and internal core may be formed from at the same time via a selective laser melting process, thus eliminating the need for using the conventional lost-wax casting system. In at least one embodiment, the external shell and internal core may be formed a ceramic material that may support receiving molten metal to form a turbine component. Once formed, the external shell and internal core may be removed to reveal the turbine component.

A component for a gas turbine engine includes a component body formed of ceramic matrix composite lamina and has at least one hook. The at least one hook has an attachment region radially inward of the at least one hook. The attachment region is radially thinner from a hook end of the at least one hook to a remote end, and then becomes radially thicker. A slot is formed through a radial thickness of the at least one hook from the hook end in a remote direction, such that there are two sections of the attachment region. A gas turbine engine is also disclosed.