A gas turbine engine defining a radial direction, an axial direction, a circumferential direction, and a longitudinal axis is provided. The gas turbine engine includes: a fan rotatable about the longitudinal axis; a turbomachine; and a housing surrounding the turbomachine and having an upper outer surface portion and a lower outer surface portion, the housing defining a first distance extending radially from the longitudinal axis to a first point located at the upper outer surface portion, the housing further defining a second distance extending radially from the longitudinal axis to a second point located at the lower outer surface portion, and the second distance is greater than the first distance.

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.

A radial inlet body of a gas turbine engine compressor is provided. The radial inlet body includes plenums arranged around opposite sides of an inlet portion defining a compressor inlet. The plenums include outer surfaces with increasing radii of curvature and a flow diverter. The flow diverter is disposed aside the inlet portion opposite a duct. The plenums are configured to direct flows from the duct around the inlet portion. The flow diverter and the plenums are configured to cooperatively encourage uniform distribution of the flows. The flow diverter includes curvilinear sides that tangentially register with and smoothly lead to the outer surfaces of the plenums and a cusp where the curvilinear sides meet. The radial inlet body further includes a front sidewall supporting the inlet portion and an aft sidewall opposite the front sidewall. The front and aft sidewalls are angled toward one another along the plenums.

A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a first upper lobe and a first lower lobe defining a first intervening recess, and a second upper lobe and a second lower lobe defining a second intervening recess, collectively defining a neck for mounting to the disk.

A turbine engine includes an engine core extending along an engine centerline and includes a compressor section, a combustor, and a turbine section in serial flow arrangement. A set of blades are circumferentially arranged in the turbine section and the compressor section. A set of dovetails mounts the set of blades to a disk, rotated about the engine centerline. Each dovetail can include a set of lobes defining complementary recesses, collectively defining a neck for mounting to the disk.

A hollow vane assembly including an open body including an interior; at least one cover support structure formed in said open body proximate the interior; a cover brazed to the open body to form at least one flow passage; and the open body and the cover configured as two piece substantially symmetrical halves.