An exhaust chamber of a steam turbine according to an embodiment includes an outer casing which includes an end wall part in an axial direction and an extension part extending upward in the axial direction from the end wall part, a first flow guide formed into an annular shape, the first flow guide forming an upstream region of a diffuser surface in a hub-side flow guide and being fixed to an upstream end portion of the extension part on a radially inner side of the diffuser surface, and a second flow guide formed into an annular shape, the second flow guide forming a downstream region of the diffuser surface at a position downstream of the first flow guide and on a radially outer side of the extension part, and being fixed to the extension part.

An exhaust chamber of a steam turbine according to an embodiment includes an outer casing which includes an end wall part in an axial direction and an extension part extending upward in the axial direction from the end wall part, a first flow guide formed into an annular shape, the first flow guide forming an upstream region of a diffuser surface in a hub-side flow guide and being fixed to an upstream end portion of the extension part on a radially inner side of the diffuser surface, and a second flow guide formed into an annular shape, the second flow guide forming a downstream region of the diffuser surface at a position downstream of the first flow guide and on a radially outer side of the extension part, and being fixed to the extension part.

A gas turbine engine for an aircraft includes a core engine assembly including a compressor section communicating air to a combustor section where the air is mixed with fuel and ignited to generate a high-energy gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section. A free turbine is configured to be driven by gas flow from the core engine. A propulsor section aft of the core engine and is driven by the free turbine. An exhaust duct routes exhaust gases from the core engine to the free turbine. The free turbine is disposed aft of the propulsor section and the exhaust duct includes an outlet aft of the propulsor section communicating gas flow to drive the free turbine. An aircraft is also disclosed.

A gas turbine engine for an aircraft includes a core engine assembly including a compressor section communicating air to a combustor section where the air is mixed with fuel and ignited to generate a high-energy gas flow that is expanded through a turbine section. The turbine section is coupled to drive the compressor section. A free turbine is configured to be driven by gas flow from the core engine. A propulsor section aft of the core engine and is driven by the free turbine. An exhaust duct routes exhaust gases from the core engine to the free turbine. The free turbine is disposed aft of the propulsor section and the exhaust duct includes an outlet aft of the propulsor section communicating gas flow to drive the free turbine. An aircraft is also disclosed.

A structure is provided that includes a perforated first skin, a second skin and a core. The core includes a first sidewall, a second sidewall, a first baffle and a second baffle. The core forms a plurality of cavities vertically between the perforated first skin and the second skin. The first baffle is connected to the perforated first skin at a first baffle first end. The first baffle is connected to the second skin at a first baffle second end by a first moveable joint. The second baffle is connected to the perforated first skin at a second baffle first end. The second baffle is connected to the second skin at a second baffle second end. A first of the cavities extends laterally between the first sidewall and the second sidewall. The first cavity extends longitudinally between the first baffle and the second baffle.

A structure is provided that includes a perforated first skin, a second skin and a core. The core includes a first sidewall, a second sidewall, a first baffle and a second baffle. The core forms a plurality of cavities vertically between the perforated first skin and the second skin. The first baffle is connected to the perforated first skin at a first baffle first end. The first baffle is connected to the second skin at a first baffle second end by a first moveable joint. The second baffle is connected to the perforated first skin at a second baffle first end. The second baffle is connected to the second skin at a second baffle second end. A first of the cavities extends laterally between the first sidewall and the second sidewall. The first cavity extends longitudinally between the first baffle and the second baffle.

An internal structure of a primary exhaust duct of a turbomachine, which has a primary wall allowing air to pass through orifices and forming an internal surface of the primary exhaust duct, an interior skin arranged inside the primary wall, and at least one separator of which a first edge region is attached to the interior skin and which has two geometries. A change from the first geometry to the second takes place when the temperature of the separator exceeds a first temperature, and the change from the second to the first takes place when the temperature of the separator drops below a second temperature. The coefficient of expansion of the separator is greater than that of the interior skin. The variation in the geometry of the separators depending on the temperature of the engine eases assembly at ambient temperature due to the compression of the separators.

An internal structure of a primary exhaust duct of a turbomachine, which has a primary wall allowing air to pass through orifices and forming an internal surface of the primary exhaust duct, an interior skin arranged inside the primary wall, and at least one separator of which a first edge region is attached to the interior skin and which has two geometries. A change from the first geometry to the second takes place when the temperature of the separator exceeds a first temperature, and the change from the second to the first takes place when the temperature of the separator drops below a second temperature. The coefficient of expansion of the separator is greater than that of the interior skin. The variation in the geometry of the separators depending on the temperature of the engine eases assembly at ambient temperature due to the compression of the separators.

An assembly is provided for an aircraft propulsion system. This assembly includes a nacelle inner structure and a deflection limiter. The nacelle inner structure includes an internal compartment and a cowl. The internal compartment is configured to house a core of a gas turbine engine. The cowl is configured to form an outer radial periphery of the internal compartment. The cowl is also configured to form an outer radial periphery of a compartment exhaust to the internal compartment at an aft end of the cowl. The deflection limiter is attached to the cowl. The deflection limiter is configured to limit radial outward movement of the cowl.

A turbine exhaust case (TEC) has an outer case and an inner case structurally interconnected by a plurality of circumferentially spaced-apart struts. At least one of the struts has an airfoil body with a hollow core. The airfoil body has opposed pressure and suction sides extending chordwise from a leading edge to a trailing edge and spanwise from a radially inner end to a radially outer end. The radially inner end of the strut has a strut wall extension that extends through the inner case to a location radially inward of the inner case relative to the central axis.