An intermediate casing includes an inner wall having an outer surface extending along a longitudinal axis, the outer surface having a first profile in a first plane, an outer wall having an inner surface a second profile in the first plane, and first through fourth arms each extending radially. The outer and inner surfaces and the first and second arms delimit a first cavity, the first cavity having a first area and the outer and inner surfaces being separated by a first distance. The outer and inner surfaces and the third and fourth arms delimit a second cavity, the second cavity having a second area and the outer and inner surfaces being separated by a second distance. The first and second profiles are such that the first area is substantially identical to the second area and the first distance is different from the second distance.

An assembly is provided for an aircraft propulsion system. This assembly includes a variable area inlet. The variable area inlet includes a moveable structure configured to move axially along an axial centerline between a first position and a second position. The variable area inlet is configured to open an airflow inlet passage into the aircraft propulsion system when the movable structure is in the first position. The variable area inlet is configured to close the airflow inlet passage when the movable structure is in the second position. The airflow inlet passage includes an inlet. The inlet extends axially along the axial centerline from a first end to a second end. A distance from the first end to the second end changes as the inlet extends laterally between a first side and a second side.

Aspects of the disclosure are directed to a seal comprising: a first leg that emanates from a center point of the seal and is configured to contact a first component, a second leg that emanates from the center point and is configured to contact a second component that is operative at a temperature that is within a range of 648 degrees Celsius to 1093 degrees Celsius, and a third leg that emanates from the center point.

An assembly is provided for an aircraft propulsion system. This assembly includes a variable area inlet. The variable area inlet includes a moveable structure configured to move axially along an axial centerline between a first position and a second position. The variable area inlet is configured to open an airflow inlet passage into the aircraft propulsion system when the movable structure is in the first position. The variable area inlet is configured to close the airflow inlet passage when the movable structure is in the second position. The airflow inlet passage includes an inlet. The inlet extends axially along the axial centerline from a first end to a second end. A distance from the first end to the second end changes as the inlet extends laterally between a first side and a second side.

A ceramic matrix composite blade includes a central core surrounded by an outer profile, the central core comprising layers of unidirectional layup having two dimension and three dimension fiber weaves; the outer profile comprises a biased weave layup that radiates toward a leading edge and a trailing edge of the blade; wherein the biased weave layup includes fibers that extend proximate an attachment region of the blade radially and axially along an airfoil portion of the blade toward the leading edge and trailing edge.

A gas turbine engine component according to an example of the present disclosure includes, among other things, an external wall extending in a thickness direction between first and second wall surfaces. The first wall surface bounds an internal cavity, and establishes at least one surface depression along an external surface contour. The external wall includes at least one cooling passage having an outlet port established along the at least one surface depression. A method of fabricating a gas turbine engine component is also disclosed.

A variable vane includes a first vane portion and a second vane portion adjacent the first vane portion. The second vane portion includes a trunnion projecting outward from a radial end of the second vane portion. The second vane portion is configured to rotate about a trunnion axis of the trunnion from a first rotational position to a second rotational position.

A variable vane includes a first vane portion and a second vane portion adjacent the first vane portion. The second vane portion includes a trunnion projecting outward from a radial end of the second vane portion. The second vane portion is configured to rotate about a trunnion axis of the trunnion from a first rotational position to a second rotational position.

An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

An apparatus for providing foreign object debris protection an air intake of an aircraft engine. The apparatus includes a frame and a plurality of cross-members. The cross-members are positioned in the frame to define a plurality of screen openings. At least one of the cross-members has an aerodynamically efficient cross section.