According to one embodiment of this disclosure a core includes a first end and a second end spaced generally opposite from the first end. The core further includes a stacking axis defined between the first end and second end and a first toroidal structure located between the first end and the second end. The first toroidal structure includes a first passage extending through the first toroidal structure in a first direction that is perpendicular to and passes through the stacking axis. The core also includes a second toroidal structure located between the first toroidal structure and the second end. The second toroidal structure includes a second passage extending through the second toroidal structure in a second direction. The first direction and the second direction are oriented along the stacking axis at a non-zero degree angle with respect to each other.

An engine component assembly includes a first engine component having a hot surface in thermal communication with a hot combustion gas flow and a cooling surface, and a second engine component having a first surface in fluid communication with a cooling fluid flow and a second surface spaced from the cooling surface to define a space. A cooling aperture extends through the second engine component. A cooling feature extends from the cooling surface of the first engine component, and is oriented relative to the cooling aperture such that the cooling fluid flow is orthogonal and non-orthogonal to different portions of the cooling feature.

An airfoil of a gas turbine engine is provided. The airfoil includes an airfoil body having at least one internal flow passage, the body having a first surface and a second surface, the first surface defining a wall of the at least one internal flow passage and a bleed port fluidly connecting the at least one internal flow passage to the second surface. The bleed port includes a bleed orifice extending from the second surface toward the internal flow passage and a bleed port cavity extending from the first surface toward the second surface, the bleed port cavity and the bleed orifice fluidly connected. The bleed port cavity is defined by a bleed port cavity wall and a base wall surrounding the bleed orifice. The bleed port cavity wall extends from the first surface to the base wall.

An assembly is provided for an aircraft propulsion system. This assembly includes a flowpath wall and an actuation system. The flowpath wall includes an inflatable bladder with a deformable face skin and an interior volume. The deformable face skin includes an exterior surface that forms a peripheral boundary of a flowpath along the flowpath wall. The interior volume extends within the inflatable bladder to the deformable face skin. The actuation system includes an air system and an actuator. The air system is fluidly coupled to the interior volume. The air system is configured to inflate or deflate the inflatable bladder to change a geometry of the exterior surface. The actuator is disposed in the interior volume. The actuator is configured to mechanically apply a force to the deformable face skin to further change the geometry of the exterior surface.

An assembly is provided for an aircraft propulsion system. This assembly includes a propulsor rotor and a flowpath wall. The propulsor rotor is rotatable about an axis. The propulsor rotor includes a plurality of propulsor blades and an inner platform. The propulsor blades are arranged circumferentially about the axis and project radially out from the inner platform. The flowpath wall is next to and downstream of the inner platform. The flowpath wall includes an inflatable bladder and a radial outer surface. The inflatable bladder is configured to change a geometry of the radial outer surface.

An assembly is provided for an aircraft propulsion system. This assembly includes a bladed rotor, an inner flowpath, an outer flowpath, a splitter and a flowpath wall. The bladed rotor is rotatable about an axis. The inner flowpath includes an inner flowpath inlet downstream of the bladed rotor. The outer flowpath includes an outer flowpath inlet downstream of the bladed rotor. The outer flowpath inlet is radially outboard of the inner flowpath inlet. The splitter is disposed radially between and partially forms the inner flowpath inlet and the outer flowpath inlet. The flowpath wall is arranged with the splitter and forms a radial inner peripheral boundary of the outer flowpath. The flowpath wall includes an inflatable bladder and a radial outer surface. The inflatable bladder is configured to change a geometry of the radial outer surface.

An assembly is provided for an aircraft propulsion system. This assembly includes a propulsor rotor and a flowpath wall. The propulsor rotor is rotatable about an axis. The propulsor rotor includes a plurality of propulsor blades and an inner platform. The propulsor blades are arranged circumferentially about the axis and project radially out from the inner platform. The flowpath wall is next to and downstream of the inner platform. The flowpath wall includes an inflatable bladder and a radial outer surface. The inflatable bladder is configured to change a geometry of the radial outer surface.

The invention relates to a turbomachine for a flight propulsion drive, comprising a core engine with a compressor, a combustion chamber, a turbine, and a heat exchanger downstream of the turbine, through which a gas flow can flow in a flow direction of the core engine, wherein, after the turbine, a flow guidance device is arranged, in order to guide the gas flow from the turbine outlet radially outward to a heat exchanger inlet, wherein the flow guidance device is arranged along the heat exchanger and defines a flow channel.